tag:blogger.com,1999:blog-26759744635248954162024-03-15T20:13:03.222-05:00An Ex Rocket Man's Take On ItThe ravings of a trained mind.Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.comBlogger472125tag:blogger.com,1999:blog-2675974463524895416.post-76254281738489368732024-03-11T11:33:00.001-05:002024-03-12T14:30:07.883-05:00More-Refined 1- vs 2-Stage to LEO<p class="MsoNormal"><o:p></o:p></p><p>Because of repeated questions from knowledgeable readers, I took a more refined look at the scenario of
chemical launch to eastward LEO at low inclination, using either an expendable two-stage to orbit
design (TSTO), or an expendable single-stage-to-orbit
design (SSTO). For this more refined
look, I added delta-vee (dV) budgets
for rendezvous and deorbit, I looked at
a more representative orbital speed requirement, and I let the second stage of a TSTO shoulder
a minority of the gravity loss (the split being arbitrary). The first stage shoulders all the drag
loss. The SSTO shoulders all of both
losses. <b>See Figure 1</b>. </p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEizPekn1B1zHCfoCWvAG_I86vUo2_Vh8snRLEYFpYPE9vgBR5VdNQ6c9NVue3BQjv6klE8gnMsnIAKKWAlWIbzBkMQCgtYtiZU-TXiKaPHNC9mfjdPHtgWqefgPkGJ59UdG5kQXlCmImeFsV4lCG8RBRgT7ikm1USCa6BqQAv_hlSHjv2BNVOE7fr9GAhyD/s1442/F1%20dV%20reqmnts.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="590" data-original-width="1442" height="164" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEizPekn1B1zHCfoCWvAG_I86vUo2_Vh8snRLEYFpYPE9vgBR5VdNQ6c9NVue3BQjv6klE8gnMsnIAKKWAlWIbzBkMQCgtYtiZU-TXiKaPHNC9mfjdPHtgWqefgPkGJ59UdG5kQXlCmImeFsV4lCG8RBRgT7ikm1USCa6BqQAv_hlSHjv2BNVOE7fr9GAhyD/w400-h164/F1%20dV%20reqmnts.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 1 – Revised dV Requirements That Are More Realistic<o:p></o:p></p><p class="MsoNormal">The 75-25 split on shouldering gravity losses is
arbitrary, but “in the ballpark”. I still picked 5% each for gravity and drag
losses, the basis being the kinetic
energy-equivalent surface circular orbit speed.
5% gravity loss would go with good kinematics off the launch pad, meaning 0.5 gee above gravity or better, or a thrust/weight of 1.5 or better at
launch. 5% drag loss would go with a
clean, slender shape, really meaning a length/diameter ratio of 6
or larger, with no steps in
diameter. <o:p></o:p></p><p class="MsoNormal"> <b><i>TSTO
Design Considerations<o:p></o:p></i></b></p><p class="MsoNormal">For the TSTO, what I
presumed was LOX-RP1 propulsion in the first stage, “compromise”-sized to improve the
ascent-averaged specific impulse (Isp),
such that the engine is just barely unseparated firing at sea
level, at 85% of max chamber pressure
Pc. I presumed LOX-LH2 propulsion in
the second stage, sized at an expansion
area ratio (A/A*) = 100, to limit engine
length. <o:p></o:p></p><p class="MsoNormal">Both the first and second stage engine technologies were
presumed to be modest technologies that do not push the state of the art (SOTA)
very hard, something that lowers
development costs that must be amortized over the launches to be made. Accordingly,
I presumed only a max Pc = 2500 psia,
and that whatever cycle it is has,
has a dumped bleed fraction of 2%.
The pressure turndown ratio (P-TDR) for throttling is only 2.5. The usual curved bell of 18-and-8-degree
profile is presumed, along with a throat
area discharge coefficient C<sub>D</sub> = 0.995. <o:p></o:p></p><p class="MsoNormal">Rather modest stage structural design technologies were also
presumed, such that both loaded-stage
inerts were 5% of stage ignition mass,
again to reduce development costs that must be amortized over the
launches to be made. 4% has been
demonstrated, but requires custom
alloys, even for expendables. The definitions are such that payload
fraction plus inert fraction plus propellant fraction sum to 1. The first stage payload is the fully-loaded
second stage, and the second stage
payload is a fixed 100 metric ton mass riding out in the open, atop the second stage. <o:p></o:p></p><p class="MsoNormal">All propulsion was initially sized for a thrust requirement
of 500,000 lb (226.76 metric tons-force,
2223.7 KN). For any ascent
engine, this was imposed at sea
level. For the TSTO second stage, this was imposed in vacuum. Performance was computed vs altitude, and those values averaged over the list of
altitudes in the altitude table. <o:p></o:p></p><p class="MsoNormal">That is not exactly correct for an “ascent-averaged
Isp”, because the vehicle does not spend
equal time at all these altitudes, but
it is well within the “ballpark”. I
compensated for any error by presuming an Isp about 2-5 s below what the sizing
calculation said. Dimensions and flow
rates depend upon sized thrust. Flow
rates and cross sectional areas scale in proportion to thrust, while linear dimensions scale in proportion
to the square root of thrust. Isp does
not scale. <o:p></o:p></p><p class="MsoNormal">The TSTO first stage sea level engine sizing to 500,000 lb
thrust is shown in Figure 2. The TSTO
second stage vacuum engine sizing to 500,000 lb thrust is shown in Figure
3. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgw8QNVuIpWUDx1NoUoPKb4-uMKj133qbCqPIpa2JZrtPP5FvRV2sLlb7k9hwkxVl5SgPmVGE_m9wvfMtfXZP62A6m6UMRUPsbtE0ZS8nxRjUov55C8Zf0IhzemCNvsNB4BAK4A6oktv8MQum9Hero3S35r1odT_m-S3k_6Sass3Lg7FIv277dMFxJIbgox/s977/F2%20underexpanded%20kerolox.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="613" data-original-width="977" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgw8QNVuIpWUDx1NoUoPKb4-uMKj133qbCqPIpa2JZrtPP5FvRV2sLlb7k9hwkxVl5SgPmVGE_m9wvfMtfXZP62A6m6UMRUPsbtE0ZS8nxRjUov55C8Zf0IhzemCNvsNB4BAK4A6oktv8MQum9Hero3S35r1odT_m-S3k_6Sass3Lg7FIv277dMFxJIbgox/w400-h251/F2%20underexpanded%20kerolox.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2 – As-Sized TSTO First-Stage Engine Data, Un-Rescaled<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgZrhWlQms9VoaOBE71Uu5KFSRmGeBQYfEZOlUMf5UL2BBi6UuR1P7el7t30BkU53sG0UeTs0Ab7Y5niuGb9bXFoSThb7np6-LJkKBAoZzMhyphenhyphen6vb8ed_StHdUSseBWWoqmBoaBuQhRL8IGvGZHPuEBSmBQ804k0qGkbXPNJmATTuA_t7C9qxGfSO3dAUKk5/s973/F3%20vacuum%20100%20LOX%20LH2.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="973" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgZrhWlQms9VoaOBE71Uu5KFSRmGeBQYfEZOlUMf5UL2BBi6UuR1P7el7t30BkU53sG0UeTs0Ab7Y5niuGb9bXFoSThb7np6-LJkKBAoZzMhyphenhyphen6vb8ed_StHdUSseBWWoqmBoaBuQhRL8IGvGZHPuEBSmBQ804k0qGkbXPNJmATTuA_t7C9qxGfSO3dAUKk5/w400-h251/F3%20vacuum%20100%20LOX%20LH2.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 3 – As-Sized TSTO Second-Stage Engine Data, Un-Rescaled<o:p></o:p></p><p class="MsoNormal"> <b><i>SSTO
Design Considerations<o:p></o:p></i></b></p><p class="MsoNormal">For the SSTO, I
looked at both LOX-LCH4 propulsion and LOX-LH2 propulsion. Such engines were “compromise”-sized for
better ascent-averaged Isp, just like
the first stage engines in the TSTO design. However,
the technology baseline presumed, pushes the SOTA very hard indeed: these designs presume a max Pc = 4000
psia, a cycle such that the dumped bleed
fraction BF = 0, and a more challenging
P-TDR = 3. (They would compare to the
SpaceX Raptor designs.) <o:p></o:p></p><p class="MsoNormal">I kept the same rather modest stage structural design
technology, with a stage inert fraction
of 5%. In this case, there is only one stage, and its 100 metric ton payload rides out in
the open, atop the stage, exactly the same as was presumed for the
TSTO. <o:p></o:p></p><p class="MsoNormal">The hydrogen-fueled version looked good enough to check the
effects of just modest-technology. That
would use the LOX-LH2 propellant ballistic models, but employ the same reduced Pc and
non-zero-BF that was used for the TSTO engine designs. The methane-fueled version had a low-enough
payload fraction to warrant skipping this look. <o:p></o:p></p><p class="MsoNormal"><b>Figure 4</b> shows the un-rescaled methane engine results
for the edge-of-the-SOTA. <b>Figure 5 </b>shows
the un-rescaled hydrogen engine results for the edge-of-the-SOTA. <b>Figure 6</b> shows an un-rescaled hydrogen
design of the same modest-technology parameters as were used in the TSTO
design. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhCNZZyYx0_6sJ-t93Y514Y0AIfa9i65clX8KPuMiqQLwedQStLvY3JhelMiLCXYB6jZZjKO9zu21OMXMMUjx0-CPHwrVDwo4j1cgrbkuPgFcOFy4EC3K7Wf6t3638_VWdvamK7zss04Gjung_TlVW_37dOPmpiIIiRwx843EpK-G92HChuWn8ooKt2TWqS/s973/F4%20underexpanded%20lox%20lch4%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="608" data-original-width="973" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhCNZZyYx0_6sJ-t93Y514Y0AIfa9i65clX8KPuMiqQLwedQStLvY3JhelMiLCXYB6jZZjKO9zu21OMXMMUjx0-CPHwrVDwo4j1cgrbkuPgFcOFy4EC3K7Wf6t3638_VWdvamK7zss04Gjung_TlVW_37dOPmpiIIiRwx843EpK-G92HChuWn8ooKt2TWqS/w400-h250/F4%20underexpanded%20lox%20lch4%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 4 – As-Sized SSTO Methane Engine, Edge-of-the-SOTA, Un-Rescaled<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjQlIyq50TND_kwY0kKv_B4sqWkBfOWLXJdiIKydPUsgD9vcyAN20FQIjOL1v6jyOSB0qj7zaiCiax9CVsfd0eb3oChrMx8NzXNXbx7_YfGvJ2f-0fFU9gNoqkVPL4LbMk81KpeGzgfzAsvBr7RKmJTlq3ESf2jgMhHa3Q53uWUgrLIgIZJNT8iVWR44XkQ/s973/F5%20underexpanded%20lox%20lh2%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="973" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjQlIyq50TND_kwY0kKv_B4sqWkBfOWLXJdiIKydPUsgD9vcyAN20FQIjOL1v6jyOSB0qj7zaiCiax9CVsfd0eb3oChrMx8NzXNXbx7_YfGvJ2f-0fFU9gNoqkVPL4LbMk81KpeGzgfzAsvBr7RKmJTlq3ESf2jgMhHa3Q53uWUgrLIgIZJNT8iVWR44XkQ/w400-h251/F5%20underexpanded%20lox%20lh2%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 5 – As-Sized Hydrogen Engine, Edge-of-the-SOTA, Un-Rescaled<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj9NaaiFCDyQ8677CVoxw1IVIfWp9CEx0W6rzY2kR46o5C1fuC9DfF3ZGgLRilDXDa6LhfYM_zjl09q023AolGdYwpuFxKfrF9D0g01K9FHJlJvxjBWydy7oJgUYzh4MeSlZbsOFuoDif5jT_cXj4hosG-7EFSY2JYJmL4s5aF329aYTEMUdmrhvbk1CXIp/s976/F6%20underexpanded%20lox%20lh2%20modest.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="609" data-original-width="976" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj9NaaiFCDyQ8677CVoxw1IVIfWp9CEx0W6rzY2kR46o5C1fuC9DfF3ZGgLRilDXDa6LhfYM_zjl09q023AolGdYwpuFxKfrF9D0g01K9FHJlJvxjBWydy7oJgUYzh4MeSlZbsOFuoDif5jT_cXj4hosG-7EFSY2JYJmL4s5aF329aYTEMUdmrhvbk1CXIp/w400-h250/F6%20underexpanded%20lox%20lh2%20modest.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 6 – As-Sized Hydrogen Engine, modest SOTA,
Un-Rescaled<o:p></o:p></p><p class="MsoNormal"> <b><i>Doing
More Detail<o:p></o:p></i></b></p><p class="MsoNormal">In my previous posting on this topic, “Launch to Low Earth Orbit: 1 Or 2 Stages?”, posted 3 March 2024, all I did was convert dV’s to mass ratios
MR, turn that into a list of mass
fractions, and then size a weight
statement from a fixed payload mass. I
used the stage ignition masses to size total thrust requirements. And that was it. <o:p></o:p></p><p class="MsoNormal">I have since added to the simple spreadsheets I used for
that analysis. If you look at the stage
overall thrust requirements and masses to be accelerated, you can choose a number of engines
appropriate for that stage, and thus
from that overall thrust requirement,
determine what those individual engine thrust ratings must be. <o:p></o:p></p><p class="MsoNormal">I created a little thrust-resize spreadsheet, which takes the as-sized engine data, and rescales them to the necessary thrust
rating. Areas and flow rates scale as
proportional to thrust, while dimensions
scale as proportional to the square root of thrust. What is important is the estimated overall dimensions
of an individual engine. Part of <b>Figure
7</b> illustrates how these engine dimensions are scaled and created from the
estimated engine sizing data. <o:p></o:p></p><p class="MsoNormal">For only a 9-engine cluster,
I worked out how to use the engine dimensions and an assumed max gimbal
angle to estimate a clearance spacing between engine bells so that gimballing
one will avoid impacting an adjacent bell.
Adding this up along a diagonal of the 9-engine cluster provides an
estimate of the min stage diameter, as
is <b>also shown in Figure 7</b>. I used
15 degrees for the max gimbal angle, an
arbitrary choice.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgYsjXDY0mC2DuUXE6aHSUJ1v7qhlEZIrABL33l60ZXIRENgRAmQm7F2iaUGaAX6Ozw2ptIv9c8115Y3_6P4PvBv-Ps4pmG-WBpmbb-Wjnd5rxdhQ8hlTtKbx9rgbj3HWvfOB1c0s3tdzYKBdC2racMEhpvzrA2cnd15ighF_2DF4KIiaIrwf8DlPs4NLko/s991/F7%20min%20D.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="626" data-original-width="991" height="253" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgYsjXDY0mC2DuUXE6aHSUJ1v7qhlEZIrABL33l60ZXIRENgRAmQm7F2iaUGaAX6Ozw2ptIv9c8115Y3_6P4PvBv-Ps4pmG-WBpmbb-Wjnd5rxdhQ8hlTtKbx9rgbj3HWvfOB1c0s3tdzYKBdC2racMEhpvzrA2cnd15ighF_2DF4KIiaIrwf8DlPs4NLko/w400-h253/F7%20min%20D.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 7 – How Engine Dimensions Determine Stage Diameter<o:p></o:p></p><p class="MsoNormal">Once you have a min stage diameter estimate, you can begin to approximate the lengths of
the tanks, engine bays, and interstages. Those lead to a vehicle length/diameter ratio
estimate, from which to judge whether
the “slender” assumption justifying lower drag loss was justified. This is based on the same diameter for the
whole vehicle, to also qualify as
“clean”, for justifying the lower drag
loss assumption. <o:p></o:p></p><p class="MsoNormal">You can use an estimate of the engine’s operating r-ratio to
split total propellant mass into oxidizer and fuel masses, in each stage. You can use the standard specific gravity
values for those propellant materials to turn those oxidizer and fuel masses
into volumes (specific gravity is numerically equal to density in metric tons
per cubic meter). Dividing volume by
base area gets you a length of the tank that is an underestimate, since there are curved pressure dome
heads. Compensate by assuming an inter-tank
length of about a diameter. <o:p></o:p></p><p class="MsoNormal">First stage (or single stage) estimated engine length is the
length of the first stage engine bay (if there is one), but is part of the overall first stage length
regardless. If there is a second
stage, there is some sort of interstage
between it and the first stage, whose
length is the estimated overall length of a second stage engine. The length of the payload is arbitrarily
assumed to be 2 diameters. <o:p></o:p></p><p class="MsoNormal">The resulting augmented spreadsheet image for the TSTO
design is shown in <b>Figure 8.</b> The
leftmost block is the original mass and thrust sizing calculations. The rest is what I added to determine engine
counts and thrusts, and to use the
re-scaled engine dimensions to do the volumes and lengths. Images of the rescaled kerosene and hydrogen
engine spreadsheets were not included,
but are reflected in the dimensional data input at top right. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEihaVD3p0NIUxltHFgTkYXCbV5Jf9KrkhdklaWunyPPs-9nLrVhtU85ffXlky1IzmmHPZ2u70R3tJloctRXDyr9eZk0z9CsiTBuOWEO8QwI1tqA_6m4dlyWZVKQRfv6mYvz2in5NNJTg8WHqKPbh_uhlek8d8DGcPYchqteBlpWs3q616CaBqkS_BtiDTEC/s1394/F8%20TSTO%20size.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="753" data-original-width="1394" height="216" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEihaVD3p0NIUxltHFgTkYXCbV5Jf9KrkhdklaWunyPPs-9nLrVhtU85ffXlky1IzmmHPZ2u70R3tJloctRXDyr9eZk0z9CsiTBuOWEO8QwI1tqA_6m4dlyWZVKQRfv6mYvz2in5NNJTg8WHqKPbh_uhlek8d8DGcPYchqteBlpWs3q616CaBqkS_BtiDTEC/w400-h216/F8%20TSTO%20size.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 8 – Spreadsheet Image For TSTO Detail Sizing<o:p></o:p></p><p class="MsoNormal">A somewhat similar-looking spreadsheet was used for the SSTO
designs, starting with the LOX-LCH4
design looked at initially in the earlier posting. That produces the detail sizing spreadsheet
image of <b>Figure 9</b>, and the
associated engine re-scale spreadsheet image of <b>Figure 10</b>. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhr02zx1SbpQfnM_huKvBRgrDo9hRVwubJVYDL0ydO3wdk1jMWBcI8YNcsRmLQyqAvi8wgNPK0ZbRS_6mPc_bpElYwMpsErABjtSE5MgzNRwWaMZxmASv9f7zM-OImc_zXlrIsmZvNIecWZCb7AAVSxapt3LotFiO_lyZHTGE38FE2cuWbIWrFdYMyWpI-v/s1325/F9%20SSTO%20CH4%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="755" data-original-width="1325" height="228" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhr02zx1SbpQfnM_huKvBRgrDo9hRVwubJVYDL0ydO3wdk1jMWBcI8YNcsRmLQyqAvi8wgNPK0ZbRS_6mPc_bpElYwMpsErABjtSE5MgzNRwWaMZxmASv9f7zM-OImc_zXlrIsmZvNIecWZCb7AAVSxapt3LotFiO_lyZHTGE38FE2cuWbIWrFdYMyWpI-v/w400-h228/F9%20SSTO%20CH4%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 9 – Spreadsheet Image For SSTO Detail Sizing, Methane,
Edge-of-SOTA<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhqVMMFc3cTUD9HEchvxBtZRSqxe-pjDO6cbf4jU2_f12PaWYwpOQEsdUkked3ynbphMja8ioxhGt6B7f4I6_rZWF1BoCJJN0cy7oz9zk5meq5R55tQCIJnmZqqcqJ-1kpu7m1bU9WUIdLI4wEB_Lgf1B2By0br0NntXXXvTf1IfebZFAAxcg8lmELW1aMH/s923/F10%20eng%20resz%20SSTO%20CH4%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="307" data-original-width="923" height="133" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhqVMMFc3cTUD9HEchvxBtZRSqxe-pjDO6cbf4jU2_f12PaWYwpOQEsdUkked3ynbphMja8ioxhGt6B7f4I6_rZWF1BoCJJN0cy7oz9zk5meq5R55tQCIJnmZqqcqJ-1kpu7m1bU9WUIdLI4wEB_Lgf1B2By0br0NntXXXvTf1IfebZFAAxcg8lmELW1aMH/w400-h133/F10%20eng%20resz%20SSTO%20CH4%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 10 – Spreadsheet Image For SSTO Engine Re-Scale, Methane,
Edge-of-SOTA<o:p></o:p></p><p class="MsoNormal"><u>The reader should be aware of one disconnect here</u>: I picked 15 engines, not 9!
The stage diameter estimate is wrong:
it is too small! That lowers the
vehicle L/D ratio even further, from the
too-low value already obtained. For this
design, the drag dV loss to cover should
have been more than the 5% used in the velocity requirements analysis shown in <b>Figure
1 above</b>. <o:p></o:p></p><p class="MsoNormal">So as it turns out,
the recommendation in the earlier posting to use the LOX-LCH4 propellant
combination for the SSTO design has been shown to be wrong! This also shows up in the 2.1% payload
fraction and the enormous 4850 metric ton ignition mass, given in <b>Figure 9 above</b>. <o:p></o:p></p><p class="MsoNormal">Accordingly, I did
another edge-of-SOTA design for the SSTO,
this time using LOX-LH2 propulsion.
The image of the detail sizing spreadsheet is given in <b>Figure 11</b>. The engine dimension re-scaling is shown in <b>Figure
12</b>. This one actually uses 9
engines, so the diameter is
“right”, and so is the L/D.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhzC59B7KUZCtKW79M4tdY5KHcRtF-pBK1bqbKa2C-IaEiBDvqcHrcfs3q3wcz3yBJP3wWdpcGRBi9n7r9HOcdIVlxf14dRSTWCAavqW_qFlcjltUtsqfws1UZPa3s9sQiEk4mP7r1IKj81PDHalBSyG3MokEbODKOyo1_v3X8xwlOPCjTZNj_JIvN23LES/s1325/F11%20SSTO%20H2%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="755" data-original-width="1325" height="228" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhzC59B7KUZCtKW79M4tdY5KHcRtF-pBK1bqbKa2C-IaEiBDvqcHrcfs3q3wcz3yBJP3wWdpcGRBi9n7r9HOcdIVlxf14dRSTWCAavqW_qFlcjltUtsqfws1UZPa3s9sQiEk4mP7r1IKj81PDHalBSyG3MokEbODKOyo1_v3X8xwlOPCjTZNj_JIvN23LES/w400-h228/F11%20SSTO%20H2%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 11 -- Spreadsheet Image For SSTO Detail Sizing, Hydrogen,
Edge-of-SOTA<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEicbTbYl_C6BMXYOiSKruJ_ijHP9CLA9_B8mk2eiWZeX2mGpvWMyn2rf55ij5nou8D0QEmINriBb47dkf5tLsjjQ9YmtrMTrSyRjrRC-Y19RhVD7itxU2KE0eZQWo63pHq9uJkX6U4XdLlrFsTbi01WebctHIgck-gMEw4spesL3-_xQ5ZNSm9F9DZDMQbN/s923/F12%20eng%20resz%20SSTO%20H%20SOTA.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="307" data-original-width="923" height="133" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEicbTbYl_C6BMXYOiSKruJ_ijHP9CLA9_B8mk2eiWZeX2mGpvWMyn2rf55ij5nou8D0QEmINriBb47dkf5tLsjjQ9YmtrMTrSyRjrRC-Y19RhVD7itxU2KE0eZQWo63pHq9uJkX6U4XdLlrFsTbi01WebctHIgck-gMEw4spesL3-_xQ5ZNSm9F9DZDMQbN/w400-h133/F12%20eng%20resz%20SSTO%20H%20SOTA.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 12 -- Spreadsheet Image For SSTO Engine
Re-Scale, Hydrogen, Edge-of-SOTA<o:p></o:p></p><p class="MsoNormal">Comparing the payload fractions and ignition masses between <b>Figures
8 and 11</b>, 7.5% and 1401 tons TSTO vs
7.5% and 1325 tons SSTO, we see pretty
much equivalent performance between the TSTO using LOX-RP1 and LOX-LH2 both at
modest engine SOTA, and the SSTO using
all-LOX-LH2, but at the edge of the
engine SOTA. Clearly the higher average
ascent Isp of the hydrogen vs the methane made a huge difference for the
SSTO, <u>more than I initially expected
to see</u>!<o:p></o:p></p><p class="MsoNormal">That brings up determining the effects of pushing the engine
SOTA so hard with the SSTO engines. To
determine that, I used the modest SOTA
hydrogen ascent engine data of <b>Figure 6 above</b>, to create yet another SSTO design
sizing, by these same methods. The detail sizing spreadsheet image is given
in <b>Figure 13</b>, with the engine
re-scale data in <b>Figure 14</b>. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh3soHqC5Tno0mNzmFInX8TyCqM7aA3y9irBpUPl00hyVK1307pCbE9IGGzIx_Y4kDIsMUWn8q0qEiY_mpD6mnrnZdpVt5H-3-4AvUh5pW5fpepgGdt11AVtQCIbyVJ3_ogD4aFhlv51oqhNn219ZFn43YgipH3bm23hsWTkD__xFem59D4kObf_uUENFUn/s1314/F13%20SSTO%20sz%20modest%20H2.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="752" data-original-width="1314" height="229" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh3soHqC5Tno0mNzmFInX8TyCqM7aA3y9irBpUPl00hyVK1307pCbE9IGGzIx_Y4kDIsMUWn8q0qEiY_mpD6mnrnZdpVt5H-3-4AvUh5pW5fpepgGdt11AVtQCIbyVJ3_ogD4aFhlv51oqhNn219ZFn43YgipH3bm23hsWTkD__xFem59D4kObf_uUENFUn/w400-h229/F13%20SSTO%20sz%20modest%20H2.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 13 -- Spreadsheet Image For SSTO Detail Sizing, Hydrogen,
Modest SOTA<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhq4mOs-KB98K7lu2wExkbEPbj6qhu4hZ1iYH0dZTevRsCX0tFB-xVOda5yYyJYDr8RsEFDADwyIqriNs7bCILiDbDJTQSrqt2YdC6UyBEm6WLOjPJl3ccXC4V_YonTdG76jNUCDDP2LSSFL0uK9zRtdxvq5s7GNueUs9x31c1KPoIud8c3f_O3NMM1KbaL/s993/F14%20modest%20H2%20ascent%20engine%20resz.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="299" data-original-width="993" height="120" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhq4mOs-KB98K7lu2wExkbEPbj6qhu4hZ1iYH0dZTevRsCX0tFB-xVOda5yYyJYDr8RsEFDADwyIqriNs7bCILiDbDJTQSrqt2YdC6UyBEm6WLOjPJl3ccXC4V_YonTdG76jNUCDDP2LSSFL0uK9zRtdxvq5s7GNueUs9x31c1KPoIud8c3f_O3NMM1KbaL/w400-h120/F14%20modest%20H2%20ascent%20engine%20resz.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 14 -- Spreadsheet Image For SSTO Engine
Re-Scale, Hydrogen, Modest SOTA<o:p></o:p></p><p class="MsoNormal">This one is not that much reduced in payload capability
(6.7% vs 7.5% for the Edge-of-SOTA SSTO and the TSTO). It increased its launch mass a little, being 1487 metric tons, vs 1325 for the edge-of-SOTA SSTO and 1401
for the TSTO. Yet they are all 3 in the
same basic class of vehicle sizes. I did
select 9 engines, so the diameter is
valid, and the L/D is “good”. <u>There is no reason the more modest
hydrogen engine technology might not serve,
and serve well</u>.<o:p></o:p></p><p class="MsoNormal"> <b><i>Results
and Conclusions<o:p></o:p></i></b></p><p class="MsoNormal" style="margin-bottom: 0in;">Sketched images for the TSTO with
modest-technology kerosene and hydrogen propulsion, the SSTO with SOTA methane propulsion, the SSTO with SOTA hydrogen propulsion, and the SSTO with modest-technology hydrogen
propulsion, are <b>given in Figures 15
through 18 below</b>, respectively. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgNAmBVHGmNMd765RphnMrqsXVHMnSdAYq3HIFsBXcm2QiAbxkGVMqGpBaRmuJd7dWpIdx1qdgQX7FQ8Cm7YBJyhXFxp63k108StRl3SdUa-LFmyIfRmoFRlnsJfHVjFdU4Yk-02kDVQRMrgdPh0mUJOLWGqiFeDVbUwKu8XJ9Vg7_YC4-iUwnNIjHZvjDt/s870/T14.5%20compare%20SSTOs.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="177" data-original-width="870" height="81" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgNAmBVHGmNMd765RphnMrqsXVHMnSdAYq3HIFsBXcm2QiAbxkGVMqGpBaRmuJd7dWpIdx1qdgQX7FQ8Cm7YBJyhXFxp63k108StRl3SdUa-LFmyIfRmoFRlnsJfHVjFdU4Yk-02kDVQRMrgdPh0mUJOLWGqiFeDVbUwKu8XJ9Vg7_YC4-iUwnNIjHZvjDt/w400-h81/T14.5%20compare%20SSTOs.png" width="400" /></a></div><p></p><p class="MsoNormal">As the <b>table above</b> indicates, <b><i>it is ascent-averaged Isp that is the
critical factor here with the SSTO</i></b>.
The big gulf between the methane and hydrogen/SOTA ascent-averaged Isp’s
corresponds to the big gulf between the payload fractions and the ignition masses. The small gap between the hydrogen/modest and
hydrogen/SOTA Isp’s corresponds to the small gap between payload fractions and
ignition masses. <o:p></o:p></p><p class="MsoNormal">Changing the propellant combination had a huge effect on
ascent-averaged Isp and the resulting sized designs. Changing how hard the hydrogen engine
technology pushes the SOTA did not have a large effect, only a smaller one. The sized design reflects exactly that. See also <b>Figure 19 below</b>. <o:p></o:p></p><p class="MsoNormal">Before I ran this more detailed design study, I thought that pushing the SOTA vs a modest
technology would have more of an effect than it actually did. Now we see:
<u>the propellant combination has the far stronger effect</u>. Go ahead and use the more modest engine
technology. That will not stop you from
doing rather well as an SSTO, as long as
you use LOX-LH2.<o:p></o:p></p><p class="MsoNormal">The hydrogen upper stage TSTO with modest engine technology
is only a little better in terms of payload fraction than the hydrogen SSTO
with modest engine technology. But, it does offer an easier path to partial
reusability, by substituting a larger
lower stage with the ability to fly back and land. That is something to consider. <o:p></o:p></p><p class="MsoNormal">The “compromise” expansion sizing approach for ascent
engines is very important, as that is
how one achieves ascent-averaged Isp values higher than an ordinary sea level
design. <o:p></o:p></p><p class="MsoNormal">That sort of “ascent-averaged Isp is dominant” outcome for
the SSTO makes me wonder if we could do better than a kerosene first stage for
the TSTO. While beyond scope here, I will look at that in a future update or
posting. The candidates are methane and
hydrogen, of course. These will be restricted to “modest engine
technology”. The same methods will be
used, as were used here. <o:p></o:p></p><p class="MsoNormal">I do expect that one or both will significantly exceed what
we can do with a modest-technology hydrogen SSTO. The problem will be the same volume issues
that afflicted the SOTA-technology methane SSTO. But we will not know, until we try. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhkuvAgNNKxp6e5wH-qb9MnhOSPHIwLRnnDsWdy8_2P15-IafJ8C0ERsvBVyaPwwYGyLzPQv8vy_e98D5r7qLI_sewk_XAJHGzFdKH-q-wt_9DWoTRKASF5qWKizUmRFImQ9fyl8xWqZ3cc3ve_6X4J-rdICZXAS5hMNo7PbJevyMlLZTyu0fnjETRb62GY/s987/F15%20TSTO%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhkuvAgNNKxp6e5wH-qb9MnhOSPHIwLRnnDsWdy8_2P15-IafJ8C0ERsvBVyaPwwYGyLzPQv8vy_e98D5r7qLI_sewk_XAJHGzFdKH-q-wt_9DWoTRKASF5qWKizUmRFImQ9fyl8xWqZ3cc3ve_6X4J-rdICZXAS5hMNo7PbJevyMlLZTyu0fnjETRb62GY/w400-h234/F15%20TSTO%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 15 – Image of Detailed Results for TSTO, Modest Kerosene and Hydrogen<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgYLm2Xft2uVeuDEOeJ4PRZ4LzO3PEEr4VHDUPosC-M3cty9QX04cTfSf1AfeBStdP41o-qncTc_MqgP7rLnlEOA2Ilp6FdVuc-BrtCAQ-6_2PXeN9FaY14_v3OTLxlPa8UumqFjJHe_q8Ke31AvXuXBJzKn6lcOq8-Zf_rb4xvMBg1Ys7SzHuMn10ThHOX/s987/F16%20SSTO%20CH4%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgYLm2Xft2uVeuDEOeJ4PRZ4LzO3PEEr4VHDUPosC-M3cty9QX04cTfSf1AfeBStdP41o-qncTc_MqgP7rLnlEOA2Ilp6FdVuc-BrtCAQ-6_2PXeN9FaY14_v3OTLxlPa8UumqFjJHe_q8Ke31AvXuXBJzKn6lcOq8-Zf_rb4xvMBg1Ys7SzHuMn10ThHOX/w400-h234/F16%20SSTO%20CH4%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 16 – Image of Detailed Results for SSTO, SOTA Methane<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiqnpCmJobqXeWfK1G1FohyphenhyphenufHTwnbcSiONEIqD8QHgfyHnbdZ-BYbrzOqisWV0UfrSENi44G7AMi8hdOS0RB2uVIPImJKmGO5Wth7HqthTrvGKb2aEbl5aN09ax5GCctLOJwc3T6CpZT4iJRPRJGY2H-wgQoCGJSxeOeL2FYXj_zv932t9J8d3Tov11RhX/s987/F17%20SSTO%20H2%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiqnpCmJobqXeWfK1G1FohyphenhyphenufHTwnbcSiONEIqD8QHgfyHnbdZ-BYbrzOqisWV0UfrSENi44G7AMi8hdOS0RB2uVIPImJKmGO5Wth7HqthTrvGKb2aEbl5aN09ax5GCctLOJwc3T6CpZT4iJRPRJGY2H-wgQoCGJSxeOeL2FYXj_zv932t9J8d3Tov11RhX/w400-h234/F17%20SSTO%20H2%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 17 – Image of Detailed Results for SSTO, SOTA Hydrogen<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjBMppe_NAZ6EvSDHyiDJeJKUN8OTbVhyX925-FNPN1Y_Nr9XOr0spDprnnpzygNRlk42x0N0KWa2pNFJ5f1c37e1dMKSuva5zMH7U9zorGwZGwqB_TeQRFxsFpUqPkb1FYLPmVz_DA4TNghOf83XsZ0rz6LhPkruXAFRBfwyzTx-NAXYP8w1GPFrbKQmgQ/s987/F18%20SSTO%20modest%20H2%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjBMppe_NAZ6EvSDHyiDJeJKUN8OTbVhyX925-FNPN1Y_Nr9XOr0spDprnnpzygNRlk42x0N0KWa2pNFJ5f1c37e1dMKSuva5zMH7U9zorGwZGwqB_TeQRFxsFpUqPkb1FYLPmVz_DA4TNghOf83XsZ0rz6LhPkruXAFRBfwyzTx-NAXYP8w1GPFrbKQmgQ/w400-h234/F18%20SSTO%20modest%20H2%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 18 – Image of Detailed Results for SSTO, Modest Hydrogen<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhA4Ytb9SB4Zfp235XEdw_0nNGSJvlViUMJj8ZX-uHTVws5sBNfd9M9bNmuEcHtDyiaisaBIP1yAAr6WPWevJR-StxcCX98KZPeiPX919wYwtN3qq0BwF05vud0Cf7tu9QbWCNn-x6EQG1ZYc2WZDDrMS168uLtM1KHAaJElxoVz4lRJsqDfAGrv_qSlfxP/s1012/F19%20compare%20SSTOs%20plots.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="342" data-original-width="1012" height="135" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhA4Ytb9SB4Zfp235XEdw_0nNGSJvlViUMJj8ZX-uHTVws5sBNfd9M9bNmuEcHtDyiaisaBIP1yAAr6WPWevJR-StxcCX98KZPeiPX919wYwtN3qq0BwF05vud0Cf7tu9QbWCNn-x6EQG1ZYc2WZDDrMS168uLtM1KHAaJElxoVz4lRJsqDfAGrv_qSlfxP/w400-h135/F19%20compare%20SSTOs%20plots.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 19 – Plots Showing Relative Effect of Engine
Technology Level and Propellant Combination<o:p></o:p></p><p class="MsoNormal">-----</p><p class="MsoNormal"><b><u>Update 3-12-2024</u>:<o:p></o:p></b></p><p class="MsoNormal">I carried out the plan outlined at the end of the article
above, to investigate two
higher-performing propellants in the TSTO.
That required sizing a LOX-LCH4 engine of modest technology to be an
ascent engine in the first stage. I
already had a LOX-LH2 ascent engine sized,
of modest technology,
investigated for the SSTO. These
were both resized to fit a 9 engine cluster of the necessary thrust, just as in the studies done in the article
above, with the updated vehicle sizing. <o:p></o:p></p><p class="MsoNormal">For these changes to the TSTO, I did not change its second stage at
all. It was, and still is,
powered by two small LOX-LH2 engines of modest technology, sized as vacuum engines with A/A* = 100, just as before. The resized modest-technology methane ascent
engine is illustrated in <b>Figure 20 below</b>. The sized TSTO vehicle with that set of modest
technology methane engines in its first stage is depicted in <b>Figures 21 and
22 below</b>. The sized TSTO vehicle with a set of modest technology hydrogen
engines in its first stage is depicted in <b>Figures 23 and 24 below</b>. <o:p></o:p></p><p class="MsoNormal">I did not see much difference between the kerosene and
methane first stage TSTO vehicles in terms of payload fraction, but the ignition weight did reduce somewhat, going to methane. A part of this is the reduced thrust
requirement reducing engine lengths, in
a vehicle whose length and diameter are primarily sensitive to engine
dimensions and number. With a hydrogen
first stage, the payload fraction
increased noticeably, and the launch
weight decreased significantly further. <o:p></o:p></p><p class="MsoNormal">I had not reduced the ascent-average Isp of the modest
technology hydrogen ascent engines by 2-5 s when I did the hydrogen TSTO in the
article above, inputting 447 s to the
vehicle sizing. Here, I did,
inputting 445 s Isp to the vehicle sizing. I ignored this small difference
making the comparison plots of trends with the two vehicles, which is <b>Figure 25 below</b>. The main takeaway is the lower slopes of the
trends with the TSTO, compared to the
steep slopes of the trends for the SSTO.
<o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal">There is a good,
simple reason for that: the TSTO
second stage is vacuum hydrogen-powered,
and shoulders the majority of the dV requirement imposed on the
vehicle. That makes the first stage mass ratios rather
small in comparison, where the added
benefit of higher first-stage Isp is “diluted” by the constant-second stage
effects. In contrast, the SSTO has to get all the dV requirement
out of its single stage. The benefits of
the higher Isp are entirely undiluted by anything, hence the effects are large, and the trend slopes are steep.</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjtNwxuzmFG__e8dKDVJA1PhFyCkoWaV7nEh2gxaRWB82yrivtho3zAjJXe3h2th6JauVLmBCw2oB4sjpDGkZ2PNAN2GEyr6I0jcMwONHNhyphenhyphenTWnMv4-YXyhcewb6NuNkAV8CAMlWIP82bw0iX2ea47-iVf35zJNd774qKdEqjgKgsqerst3UmdMktzSWW34/s974/F20%20underexpanded%20lox%20lch4%20modest.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="607" data-original-width="974" height="249" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjtNwxuzmFG__e8dKDVJA1PhFyCkoWaV7nEh2gxaRWB82yrivtho3zAjJXe3h2th6JauVLmBCw2oB4sjpDGkZ2PNAN2GEyr6I0jcMwONHNhyphenhyphenTWnMv4-YXyhcewb6NuNkAV8CAMlWIP82bw0iX2ea47-iVf35zJNd774qKdEqjgKgsqerst3UmdMktzSWW34/w400-h249/F20%20underexpanded%20lox%20lch4%20modest.png" width="400" /></a></div><p class="MsoNormal">Figure 20 – Sized Methane Ascent Engine of Modest Technology<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjlWQm7nIZ5hIGhv2cr6p64BVkl4J7bCDsWep8fmXF_4_AacX8_1JdUYE5U3hpKE3MmOEelqlIPapJ1txLUbGiHuTjBEC8t93Z0_bFXjn2oz-KgEgSTEaVIffmNEnKzdfERHAhDR0tBX32TDtfcLU9lBNr9k8JrSpvJQRV81vqmnK1vjJi7zeTdHbiqJc-K/s1394/F21%20TSTO%20modest%20methane.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="747" data-original-width="1394" height="214" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjlWQm7nIZ5hIGhv2cr6p64BVkl4J7bCDsWep8fmXF_4_AacX8_1JdUYE5U3hpKE3MmOEelqlIPapJ1txLUbGiHuTjBEC8t93Z0_bFXjn2oz-KgEgSTEaVIffmNEnKzdfERHAhDR0tBX32TDtfcLU9lBNr9k8JrSpvJQRV81vqmnK1vjJi7zeTdHbiqJc-K/w400-h214/F21%20TSTO%20modest%20methane.png" width="400" /></a></div><p class="MsoNormal">Figure 21 – Vehicle Sizing Data for Modest-Technology
Methane Engines in the First Stage<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiGakJsGNRUQG-QhpLig9Ptp71wkbG8lrA_UlvjeiSdjy5DmUs2cObiOww1j-DrrZxL5FkGt9_RYp6D5selXSb3B3JEaYuoRvfBMrqUtZF1bQidcMtZUAS-H1brQekjka0b8FS9H9RPnSbBSat_9rbBKl52BLZkhfKZx8exoUdS-4lG8iZkvmystHHRYDhx/s987/F22%20TSTO%20modest%20methane%201st.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiGakJsGNRUQG-QhpLig9Ptp71wkbG8lrA_UlvjeiSdjy5DmUs2cObiOww1j-DrrZxL5FkGt9_RYp6D5selXSb3B3JEaYuoRvfBMrqUtZF1bQidcMtZUAS-H1brQekjka0b8FS9H9RPnSbBSat_9rbBKl52BLZkhfKZx8exoUdS-4lG8iZkvmystHHRYDhx/w400-h234/F22%20TSTO%20modest%20methane%201st.png" width="400" /></a></div><p class="MsoNormal">Figure 22 – Vehicle Sketch for Modest-Technology Methane
Engines in the First Stage<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjqPcDufGTBLfyVo4If7R5oL6eVW8d4ly8-P1mIAcqxvRrK3YU9_n2-B90XC_TskS88ML2l2EVoGRQcq6yT7uoWQ0aON9kHbU4TMpAuxNvKOQU4VEPuv0pWJX3mhRL5hWXgHrRN2dcHWRGhS6OcBq9ukvMGBfEY9S4WHzzwrTSrwP-ygQJ8RRkgTZIUBF1Q/s1387/F23%20TSTO%20modest%20hydrogen.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="741" data-original-width="1387" height="214" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjqPcDufGTBLfyVo4If7R5oL6eVW8d4ly8-P1mIAcqxvRrK3YU9_n2-B90XC_TskS88ML2l2EVoGRQcq6yT7uoWQ0aON9kHbU4TMpAuxNvKOQU4VEPuv0pWJX3mhRL5hWXgHrRN2dcHWRGhS6OcBq9ukvMGBfEY9S4WHzzwrTSrwP-ygQJ8RRkgTZIUBF1Q/w400-h214/F23%20TSTO%20modest%20hydrogen.png" width="400" /></a></div><p class="MsoNormal">Figure 23 – Vehicle Sizing Data for Modest-Technology
Hydrogen Engines in the First Stage<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi0GqjGmxM2Mmmsyqi5P2ujqWfUUF2u-yPJkiu1EqhdBRK7JANllF5jtqIwTkDIhiQVZT80GLxLgl3i48Rd-oGaySTCReMNApLlP8wEftcEkPYAIwr8LIquqWkFSH-wMHXr5McGdvFmDE9VMP24vVckYuCigd3qUJK563iLKePzZpYpeUxA0tZMBwByF5TH/s987/F24%20TSTO%20modest%20hydrogen%201st.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="578" data-original-width="987" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi0GqjGmxM2Mmmsyqi5P2ujqWfUUF2u-yPJkiu1EqhdBRK7JANllF5jtqIwTkDIhiQVZT80GLxLgl3i48Rd-oGaySTCReMNApLlP8wEftcEkPYAIwr8LIquqWkFSH-wMHXr5McGdvFmDE9VMP24vVckYuCigd3qUJK563iLKePzZpYpeUxA0tZMBwByF5TH/w400-h234/F24%20TSTO%20modest%20hydrogen%201st.png" width="400" /></a></div><p class="MsoNormal">Figure 24 – Vehicle Sketch for Modest-Technology Hydrogen
Engines in the First Stage<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhprpc24g-GwS_q95vrYbdGVKvQkWj9aJx5INC_BBxGsxIHZMXlwvBeH1NVeHXtpsOt_YjkRLTEmhWh1djGMlXAXkrZskbCC3sKknYp8_jPcWHcKyUxfkw59nbgNiv5o58g2V4HGSjnv-BKlb2qhVggG-f66ljQwjAz-k7aSPp5l3r2n9DjVdu_TFYT50CC/s1002/F25%20all%20compare.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="588" data-original-width="1002" height="235" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhprpc24g-GwS_q95vrYbdGVKvQkWj9aJx5INC_BBxGsxIHZMXlwvBeH1NVeHXtpsOt_YjkRLTEmhWh1djGMlXAXkrZskbCC3sKknYp8_jPcWHcKyUxfkw59nbgNiv5o58g2V4HGSjnv-BKlb2qhVggG-f66ljQwjAz-k7aSPp5l3r2n9DjVdu_TFYT50CC/w400-h235/F25%20all%20compare.png" width="400" /></a></div><p class="MsoNormal">Figure 25 – Comparison Plots of Trends, With All Vehicles<o:p></o:p></p>
<p class="MsoNormal"> <b><i>Conclusions<o:p></o:p></i></b></p>
<p class="MsoNormal">The results here are for all-expendable vehicle
sizings. The conclusions apply to the
same, with exceptions for re-usability as
stated in notes 7 and 8. <o:p></o:p></p>
<p class="MsoNormal">#1. If you design a TSTO expendable “from scratch” for
delivering large payloads to LEO, always
use a LOX-LH2 engine designed for vacuum operation to power the second
stage. <o:p></o:p></p>
<p class="MsoNormal">#2. If you design a TSTO expendable “from scratch” for
delivering large payloads to LEO, it
does not matter very much which of the 3 propellant combinations you use for
powering the first stage. The trends
favor LOX-LH2, but these trends are weak
(low slope). LOX-RP1 and LOX-CH4 also
serve well.<o:p></o:p></p>
<p class="MsoNormal">#3. Whether you design “from scratch” a TSTO expendable or
an SSTO expendable for delivering large payloads to LEO, use “ascent engines” with their expansion ratio
designed as an ascent compromise: just
barely unseparated, at around 85% max Pc,
at sea level. Engines designed in this way will have a
higher ascent-averaged Isp than traditional sea level engine designs, which are generally perfectly expanded to sea
level pressure at max Pc. And the actual
flight configurations are testable at
sea level in the open-air nozzle mode,
which helps to greatly lower development costs that must be
amortized, and to greatly lower
development risks. <o:p></o:p></p>
<p class="MsoNormal">#4. If you design “from scratch” an SSTO expendable for
delivering large payloads to LEO, go for
the LOX-LH2 propulsion. Because of the
steep trends, these designs are critically-sensitive
to ascent-averaged Isp above all other considerations. Only LOX-LH2 provides high enough Isp. <o:p></o:p></p>
<p class="MsoNormal">#5. Neither type of vehicle is extremely-sensitive to how
hard the engine technology pushes the SOTA,
because the Isp difference is not all that large between high-SOTA and
rather modest technology, something true
for all 3 propellant combinations. With
the more modest technology, development
risks and efforts are lower, leading to
lower development costs to be amortized. <o:p></o:p></p>
<p class="MsoNormal">#6. I did <u>not</u> evaluate the impact of stage structural
design technology! I got good results
from the best of the designs at a rather modest stage inert mass fraction
assumption: 5% inert in every loaded
stage. 4% has been demonstrated, but I deliberately chose not to push those
limits! The less demanding structural
design lowers development effort levels and development risks, thus lowering the development costs to be
amortized. <o:p></o:p></p>
<p class="MsoNormal">#7. The TSTO offers a fairly easy path to partial
re-usability, by enlarging the first
stage design to enable its flyback,
entry, and recovery. This is primarily enabled by the relatively-low
(only supersonic) speeds at entry, in
turn imposed by the relatively low staging speed, which also lowers the burn-back dV
requirement.<o:p></o:p></p>
<p class="MsoNormal">#8. The SSTO does <u>not</u> offer an easy path to
re-usability, because the entry speeds
are orbital-class hypersonic, and the
stage simply does not have the inert fraction to permit the design changes to
make it into a survivable entry vehicle at all,
much less to land. The “proof” is
in the negative: if this were not true
as stated, it would have already been
done, routinely, along with first stage recoveries. <o:p></o:p></p>
<p class="MsoNormal"> <b><i>Final
remarks<o:p></o:p></i></b></p>
<p class="MsoNormal">Do not take these “designs” as ready-to-build! While the engine ballistics and performance
estimates are rather good, the weight
statements are less so, and the
dimensional estimates are only “ballpark”.
It is the <u>trends that should be used</u> to support real design
candidate screening and selections. Some
of that screening I have done for you,
in this article. <o:p></o:p></p>
<p class="MsoNormal">To address questions from knowledgeable readers, I made the dV requirements more
representative of vehicles that can get to orbit and rendezvous with a
destination, plus a deorbit capability
for proper disposal. But, there are a lot of things that I did not
address.<o:p></o:p></p>
<p class="MsoNormal">I did <u>not</u> address propellant ullage / engine relight
issues, and I did <u>not</u> address the
unrecoverable propellant fractions that are inherent with any type of tank
design. Further, I did <u>not</u> address the actual end dome
shapes or designs of the liquid propellant tanks, or the possibility of a common dome
design, which can be done with some
propellant combinations, but by no means
all of them. <o:p></o:p></p>
<p class="MsoNormal">These are not only “from scratch” vehicle ballpark design
sizings, they are also “from scratch”
paper engine design sizings, a
start-point only for a real engine design and development effort. I made absolutely no attempt in this
work, to use any pre-existing engine
designs of any kind at all! <o:p></o:p></p>
<p class="MsoNormal">My work here can be re-scaled to other delivered payload
masses (the 100 metric tons that I used here was an arbitrary number), so that the trends I uncovered can help guide
real concept selection and real design efforts for other-size projects done by
others. If a pre-existing engine of the
right propellant combination fits your design project, so much the better! Any development costs you can avoid are one
less thing to amortize over the life of the product. <o:p></o:p></p><p class="MsoNormal"><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-10174998533172516582024-03-04T08:23:00.003-06:002024-03-05T17:52:43.324-06:00Launch to Low Earth Orbit: Fixed-Geometry Options<p>The question keeps coming up among enthusiasts about how
fixed-geometry conventional bell-type rocket engines cannot adequately serve
for ascent through the atmosphere. This
“justifies” adding technologies such as bell extensions or free-expansion
nozzle designs. This document addresses
what can be done with a fixed-geometry expansion bell, if you carefully design it for that ascent purpose, and subject to the constraint that you can
test the flight article at sea level,
open-air nozzle conditions, for
both cost-effectiveness, and for testing
the actual flight article. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">For this study,<span style="mso-spacerun: yes;"> </span>I
presumed liquid oxygen/liquid hydrogen (LOX-LH2) propellants,<span style="mso-spacerun: yes;"> </span>and a modest engine design technology
characterized as max chamber pressure (Pc) 2500 psia,<span style="mso-spacerun: yes;"> </span>with a pressure turndown ratio (P-TDR) of
2.5.<span style="mso-spacerun: yes;"> </span>Expansion bells are curved,<span style="mso-spacerun: yes;"> </span>with an “18-8” degree profile,<span style="mso-spacerun: yes;"> </span>for a constant nozzle kinetic energy efficiency
(η<sub>KE</sub>) just over 98.7%.<span style="mso-spacerun: yes;"> </span>Throat
discharge coefficient was also constant at C<sub>D</sub> = 0.995.<span style="mso-spacerun: yes;"> </span>The engine cycle is modeled as “modest modern
technology” with a bleed dump fraction (BF) of 2%.<span style="mso-spacerun: yes;"> </span>All these variables were held exactly the
same,<span style="mso-spacerun: yes;"> </span>for the 3 designs explored
here,<span style="mso-spacerun: yes;"> </span>to investigate the effects of
expansion and thrust sizing effects.<o:p></o:p></p>
<p class="MsoNormal">I explored 3 designs:<span style="mso-spacerun: yes;">
</span>(1) a conventional “sea level-optimized” design,<span style="mso-spacerun: yes;"> </span>(2) a typical conventional “vacuum optimized”
design,<span style="mso-spacerun: yes;"> </span>and (3) the practical compromise
design that I favor for best ascent performance.<span style="mso-spacerun: yes;"> </span>The sea level “optimization” requires that
the bell expansion be sized for max chamber pressure expanding down to standard
sea level pressure,<span style="mso-spacerun: yes;"> </span>and with throat and
exit dimensions (and total flow rate) sized so that sea level thrust at max Pc
is a required value.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">There is really no such thing as a “vacuum-optimized”
expansion bell!<span style="mso-spacerun: yes;"> </span>Such would have an
infinite exit area expanding the flow to exactly zero exit plane pressure,<span style="mso-spacerun: yes;"> </span>which is utter nonsense.<span style="mso-spacerun: yes;"> </span>Vacuum engine bell designs are not optimized
in any way,<span style="mso-spacerun: yes;"> </span>they are merely constrained
by practical design constraints.<span style="mso-spacerun: yes;"> </span>The
word “optimized”,<span style="mso-spacerun: yes;"> </span>while routinely
used,<span style="mso-spacerun: yes;"> </span><u>is a complete misnomer</u> for
such things. <o:p></o:p></p>
<p class="MsoNormal">The length and diameter of the bell has to fit practical
dimensional constraints so that the engine will fit within the available space
at the rear end of the stage,<span style="mso-spacerun: yes;"> </span>often even
more severely constrained than one might think,<span style="mso-spacerun: yes;">
</span>if the engine must gimbal for thrust vectoring.<span style="mso-spacerun: yes;"> </span>For purposes of this study,<span style="mso-spacerun: yes;"> </span>I simply analyzed to a convenient “typical”
approximation:<span style="mso-spacerun: yes;"> </span>a nozzle area expansion
ratio of 100.<span style="mso-spacerun: yes;"> </span>These cannot be fired
open-air nozzle at sea level,<span style="mso-spacerun: yes;"> </span>even at
full throttle.<span style="mso-spacerun: yes;"> </span>Higher Pc lowers the
separation altitude,<span style="mso-spacerun: yes;"> </span>but does not
eliminate the sea level separation problem.<o:p></o:p></p>
<p class="MsoNormal">My preferred sizing approach defines a high part-throttle
setting at which the bell expands from the part-throttle Pc to an over-expanded exit
plane pressure Pe that corresponds to being on the verge of
backpressure-induced flow separation at sea level.<span style="mso-spacerun: yes;"> </span>This allows sea level open-air nozzle testing
of the full flight configuration at that part-throttle setting or higher,<span style="mso-spacerun: yes;"> </span>a considerable cost savings in development
testing!<span style="mso-spacerun: yes;"> </span>Such designs are over-expanded, <span style="mso-spacerun: yes;"> </span>and thus underperform the sea level designs at
sea level.<span style="mso-spacerun: yes;"> </span>But,<span style="mso-spacerun: yes;"> </span>they also offer considerably more expansion
ratio,<span style="mso-spacerun: yes;"> </span>and so perform nearly as well as
the so-called “vacuum optimized” designs,<span style="mso-spacerun: yes;">
</span>out in vacuum.<span style="mso-spacerun: yes;"> </span>It’s a compromise
for ascent!<o:p></o:p></p>
<p class="MsoNormal">For such a compromise design to serve in ascent,<span style="mso-spacerun: yes;"> </span>I size its dimensions and flow rate to meet
the imposed thrust requirement at sea level.<span style="mso-spacerun: yes;">
</span>(For “vacuum optimized” designs that can still be tested at sea
level,<span style="mso-spacerun: yes;"> </span>I size the dimensions and flow
rates to a thrust requirement imposed out in vacuum;<span style="mso-spacerun: yes;"> </span>that was not done for this study.)<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">The normal approach to “vacuum-optimized” designs results in
larger expansion ratios that simply cannot be tested at sea level open-air
nozzle at all!<span style="mso-spacerun: yes;"> </span>The only feasible way to
test them open-air at sea level,<span style="mso-spacerun: yes;"> </span>is to
fit them with shorter test bells that do not flow-separate.<span style="mso-spacerun: yes;"> </span>You cannot surface-test the full flight
article that way,<span style="mso-spacerun: yes;"> </span>but that is what is
often done,<span style="mso-spacerun: yes;"> </span>anyway.<o:p></o:p></p>
<p class="MsoNormal"><span style="mso-tab-count: 1;"> </span><b><i>Sea
Level Design<o:p></o:p></i></b></p>
<p class="MsoNormal">The sea level design is depicted in <b>Figure 1</b>.<span style="mso-spacerun: yes;"> </span>Expansion was from max Pc to Pe = sea level
standard pressure.<span style="mso-spacerun: yes;"> </span>Thrust was sized to a
500,000 lb (226.7 metric tons-force,<span style="mso-spacerun: yes;">
</span>2223.7 KN) requirement imposed at sea level,<span style="mso-spacerun: yes;"> </span>full throttle conditions.<span style="mso-spacerun: yes;"> </span>I used my “compressible.xlsx” spreadsheet
tool for this effort,<span style="mso-spacerun: yes;"> </span>specifically the
“r noz alt” worksheet so that performance versus altitude could be
plotted,<span style="mso-spacerun: yes;"> </span>and ascent-averaged.<span style="mso-spacerun: yes;"> </span>Selected items from the spreadsheet were
copied to the figure,<span style="mso-spacerun: yes;"> </span>along with the two
plots the worksheet creates.<span style="mso-spacerun: yes;"> </span>These were
annotated as shown.<span style="mso-spacerun: yes;"> </span>Note how the sea
level design can be tested open-air nozzle at sea level,<span style="mso-spacerun: yes;"> </span>even at the minimum throttle setting.<span style="mso-spacerun: yes;"> </span>Thrust in vacuum is only a little higher than
sea level.<span style="mso-spacerun: yes;"> </span>Vacuum specific impulse (Isp)
is much higher than the sea level value,<span style="mso-spacerun: yes;">
</span>but the ascent-averaged Isp is only a little lower than the vacuum
value. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEij36KQO7RlDU7pBWThW-xDdAwpamXEDTvHcE2KSu2akqbEBjUsTkCcxe4fNdNi70oLzbx9u1EQO3v6Rd3Bv0cqpX8h7R_L2bagAnksbAOlpG-rPo0Qu8GmD7A5ZoG7e0Y7ZDXAgPHiyGVpVFlgk0nG8lep5SqvRvtvP0RegPZ7hguNp_ehTT7k51V0T4GB/s974/F1%20sea%20level.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="609" data-original-width="974" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEij36KQO7RlDU7pBWThW-xDdAwpamXEDTvHcE2KSu2akqbEBjUsTkCcxe4fNdNi70oLzbx9u1EQO3v6Rd3Bv0cqpX8h7R_L2bagAnksbAOlpG-rPo0Qu8GmD7A5ZoG7e0Y7ZDXAgPHiyGVpVFlgk0nG8lep5SqvRvtvP0RegPZ7hguNp_ehTT7k51V0T4GB/w400-h250/F1%20sea%20level.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Standard Sea Level Design, As Is Often Used In First Stages for Ascent<o:p></o:p></p>
<p class="MsoNormal"> <b><i>Typical
of Standard “Vacuum-Optimized” Designs<o:p></o:p></i></b></p>
<p class="MsoNormal">This design is depicted in a very similar format in <b>Figure
2</b>. Expansion was sized from max Pc
to a Pe that produced A/A* = 100.
Dimensions and flow rates were sized from a thrust requirement imposed
in vacuum. This was the same 500,000 lb
(226.7 metric tons-force, 2223.7
KN). The vacuum performance is quite
good, as one would expect. This design cannot be used for ascent, or tested open-air nozzle at sea level, because it suffers backpressure-induced flow
separation at all throttle settings. The
nozzle is separated below about 20 kft at 100% Pc, separated below about 25 kft at 85% Pc, and separated below about 40 kft at the min
setting of 40% Pc. The as-sized
dimensions and flow rate are different from those of the sea level design, but not very much different at all! A common chamber power head could be used
with two different bells, by just
adjusting one thrust requirement a little bit.
That was not done here, though.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgqUTgYppQDwXOTwpLOIvxj2WMJvELh2YMDok4Su1S47aUNkwCueu47Bl3SFuzLiflpUed-Ovk5zoPwt_uvNBtrh58Zj6-dg1nYO1DpWsO8lGxGXx7zoQAlQRYVXbVXPgZMOzUkmhbsJZs4RdvH3O69Tr4oRwZFlvyHQ4Y7qEOHAKJOXQufCycnMxoF0sQm/s976/F2%20vacuum.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="976" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgqUTgYppQDwXOTwpLOIvxj2WMJvELh2YMDok4Su1S47aUNkwCueu47Bl3SFuzLiflpUed-Ovk5zoPwt_uvNBtrh58Zj6-dg1nYO1DpWsO8lGxGXx7zoQAlQRYVXbVXPgZMOzUkmhbsJZs4RdvH3O69Tr4oRwZFlvyHQ4Y7qEOHAKJOXQufCycnMxoF0sQm/w400-h250/F2%20vacuum.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Typical of Standard “Vacuum-Optimized”
Designs, Cannot Be Used for Ascent<o:p></o:p></p>
<p class="MsoNormal"> <b><i>Compromise
Design Intended For Ascent Use<o:p></o:p></i></b></p>
<p class="MsoNormal">This design is depicted in <b>Figure 3</b>, using a format very similar to the other two
designs. The expansion was sized for 85%
Pc down to a Pe = 3.63 psia that had just barely above standard sea level
pressure as its flow separation backpressure.
This is overexpanded, as much as
can be tolerated at that throttle setting,
without actually separating. That produces a bigger area expansion ratio
than a sea level design, while still
less than most “vacuum-optimized” designs.
<o:p></o:p></p>
<p class="MsoNormal">Being intended for ascent use, I imposed the thrust requirement at sea
level, as the same 500,000 lb (226.7
metric tons-force, 2223.7 KN). Both this and the standard sea level design
meet this thrust requirement at sea level and 100% throttle. The compromise design thrusts better out in
vacuum than the sea level design, at
about 587,000 lb, versus only about
531,000 lb. The “vacuum-optimized”
design delivers 500,000 lb in vacuum,
exactly per its design. <o:p></o:p></p>
<p class="MsoNormal">The real story is in terms of Isp. The compromise design is overexpanded and has
lower Isp at sea level than the sea level design. However,
its vacuum performance is only a little bit less than that of the so-called
“vacuum-optimized “ design, which cannot
be used at sea level at all!
Further, its ascent-averaged Isp
is actually greater than that of the sea level design. The compromise design can be used for ascent
from sea level to vacuum, with better
ascent-averaged Isp than the sea level design, and nearly the vacuum Isp of the
“vacuum-optimized” design! And it does
this with the same high thrust/weight ratio,
and fewer potential failure modes,
as are inherent in all fixed geometry bell designs. Further,
it is testable at sea level in open-air nozzle mode, and in the full flight configuration, a real plus!<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjP9mTHDfFR9o51FHBMIRMF50jG_vLopwCvsUs5sfFZIn_497JGSXv30snoKd55D8faBApcA-obKNdqAx2l_bsxafEpKiwp63GQCoohpGX4aFG9qzIOjHcsH6u-Y9mm0t2jBb6HKM0wqj0D6MJN4wkT962iNiy_HKTNpAWWQbh28KYff8dXU0qhDA2dELdg/s974/F3%20underexpanded.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="609" data-original-width="974" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjP9mTHDfFR9o51FHBMIRMF50jG_vLopwCvsUs5sfFZIn_497JGSXv30snoKd55D8faBApcA-obKNdqAx2l_bsxafEpKiwp63GQCoohpGX4aFG9qzIOjHcsH6u-Y9mm0t2jBb6HKM0wqj0D6MJN4wkT962iNiy_HKTNpAWWQbh28KYff8dXU0qhDA2dELdg/w400-h250/F3%20underexpanded.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – Compromise Design Intended For Ascent Use<o:p></o:p></p>
<p class="MsoNormal"> <b><i>Summary
Comparison<o:p></o:p></i></b></p>
<p class="MsoNormal">The table shows a summary comparison of the 3 designs. All this was in the figures in more detail.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiKp7E_krRK8Z2mEMpdcBCoJ55ahNFqxRGYKvEy_Bf1WRltWK9dO9Xi7KS4y2H7aRqxeqQ6Zpfi3QP1D9rG_jy736ix1pBjzEXjb2xO8YSYbqPVvW0dojAeQeZlTnXYDEILeli6WSiNz7H40w62GuDliJZu-qjMQye9mQxk7a7jSxhuEnAMTLpEqXmlajiV/s699/table.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="306" data-original-width="699" height="175" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiKp7E_krRK8Z2mEMpdcBCoJ55ahNFqxRGYKvEy_Bf1WRltWK9dO9Xi7KS4y2H7aRqxeqQ6Zpfi3QP1D9rG_jy736ix1pBjzEXjb2xO8YSYbqPVvW0dojAeQeZlTnXYDEILeli6WSiNz7H40w62GuDliJZu-qjMQye9mQxk7a7jSxhuEnAMTLpEqXmlajiV/w400-h175/table.png" width="400" /></a></div><p class="MsoNormal" style="text-indent: 0.5in;"><b><i>Discussion of Other Design
Alternatives<o:p></o:p></i></b></p>
<p class="MsoNormal">There are two design alternatives to the fixed-geometry
expansion bell: (1) the extendible bell
extension, and (2) the free-expansion
approach. <o:p></o:p></p>
<p class="MsoNormal">The <b><span style="background: yellow; mso-highlight: yellow;">extendible
bell extension</span></b> is depicted in <b>Figure 4</b>, which I got from R. Gregory Clark’s
“Polymath” site “exoscentist.blogspot.com”.
Basically, the idea is to add the
bell extension hardware to a sea level design,
which converts it to a “vacuum-optimized” design geometry with the same
chamber power head. You would extend the
bell section somewhere above the altitude at which the vacuum geometry risks
flow separation, something subject to
some design optimization, of course. But,
that would always occur somewhere in the lower stratosphere, in any event,
as can be seen by the curves in <b>Figure 2 above</b>, which have a definite “knee” somewhere near
50 kft altitude. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgD2C1p-MEWQ9E2CoK2wyIviF54tb5x0A5wQOvAPMzwhP08KVb9USrW0lJj36w9FmxGz7E4RF-af5nc39bXhwmQPPBfvRMH-aV_9u5TW3410MLm1M3SIelRvbjCfh7kKkfdCmbHYAz7BZ68MHklvrXuqSO49wEpJbdpWqYY5ST9IlPBMvHdJupAesO6GH5_/s1157/F4%20extend%20bell.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1007" data-original-width="1157" height="349" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgD2C1p-MEWQ9E2CoK2wyIviF54tb5x0A5wQOvAPMzwhP08KVb9USrW0lJj36w9FmxGz7E4RF-af5nc39bXhwmQPPBfvRMH-aV_9u5TW3410MLm1M3SIelRvbjCfh7kKkfdCmbHYAz7BZ68MHklvrXuqSO49wEpJbdpWqYY5ST9IlPBMvHdJupAesO6GH5_/w400-h349/F4%20extend%20bell.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Depiction of the Extendible Bell Extension
Concept – From R G Clark’s “Polymath” Site<o:p></o:p></p>
<p class="MsoNormal">This approach will indeed increase the ascent-averaged
Isp, to something between the vacuum
design value and the sea level design value,
and likely much closer to the vacuum value, as we have already seen. But, besides
imposing geometric restrictions upon the shape and placement of the chamber
power head and plumbing items, <b><i>adding
this kind of hardware is going to increase engine inert mass! </i></b> <o:p></o:p></p>
<p class="MsoNormal">You will see this as a substantially-lower engine
thrust/weight ratio compared to simpler fixed-geometry designs. That will adversely-affect stage or vehicle
inert masses, and thus also mass ratio
and velocity-increment capabilities. You
will not see as much (if any) improvement over the compromise design, as otherwise only the ascent-averaged Isp
might suggest. <o:p></o:p></p>
<p class="MsoNormal">Further, adding
moving hardware with sealing issues <b><i>adds a long list of potential failure
modes</i></b> to the list of potential failure modes that any rocket engine
has. This is a safety/reliability
thing, something not so easily
quantified, but very real, nonetheless!
I have no numbers to show, <u>but
I do have to ask the very pertinent question</u>: <b><i>why would you do this, unless you were <u>absolutely forced</u> into
taking the increased risks?<o:p></o:p></i></b></p>
<p class="MsoNormal">The <b><span style="background: yellow; mso-highlight: yellow;">free-expansion
nozzle design approach</span></b> is something I have explored in other
articles on this site. It can take many
forms, but the best-known is the
aerospike nozzle approach. This can be a
coaxial or a 2-D linear design. The 2-D
linear design was intended to be implemented on the X-30 “Venture Star” design,
for one-stage low orbit access, that ultimately proved unsuccessful for a
variety of reasons. Most of those had to
do with losing strength by reducing inert weight too far.<o:p></o:p></p>
<p class="MsoNormal"><b><i>These free-expansion designs all suffer from a common
problem: as the ambient atmospheric
pressure drops to low or zero values,
the streamlines of the propulsion stream diverge excessively and very
adversely, in terms of critically-low effective
kinetic energy efficiency.</i></b> This
drastically reduces specific impulse (Isp) at high altitudes and on out into
vacuum, and it is inherent, being fundamental compressible flow physics! <b><i>These things work better than
conventional nozzles from the surface out to the lower stratosphere, but from there on up, Isp performance falls very drastically out
into vacuum. </i></b>They are absolutely
lousy as vacuum engines, and they inherently
always will be, despite the common false
perceptions to the contrary! <o:p></o:p></p>
<p class="MsoNormal">With the fixed-bell designs showing ascent-averaged Isp
closer to the vacuum value than the sea level value, it is to be reasonably expected that
aerospike ascent-averaged Isp will be closer to the utterly-lousy vacuum Isp
levels that these designs produce. Even
if not, the huge deficit the lousy
vacuum performance represents, will drag
the ascent-averaged Isp down catastrophically,
<u>no matter how it is computed</u>. <o:p></o:p></p>
<p class="MsoNormal">The earlier articles on this site that explore
free-expansion nozzle performance are listed below. Use the blog archive gadget on the left of this
page to find them very quickly. <u>All
you need is the title and the date</u>.
Click on the year, then the
month, then the title if more than one
article was posted that month. <o:p></o:p></p>
<p class="MsoNormal">Rocket Nozzle Types (bells and aerospikes) 4 February 2023<o:p></o:p></p>
<p class="MsoNormal">How Propulsion Nozzles Work
12
November 2018<o:p></o:p></p><p class="MsoNormal"><b>Addendum<o:p></o:p></b></p><p class="MsoNormal">The “compressible.xlsx” spreadsheet was put together to
support a course I created in the basics of compressible flow applications. Rocket nozzles are but one application of
this. There are multiple worksheets in
that spreadsheet file, of which only three
relate to estimating rocket engine performance.
<o:p></o:p></p><p class="MsoNormal">One (“prop comb”) is just a data library of supporting
ballistics data for several propellant combinations. Another one (“rocket noz”) does the basic
sizing calculations with sea level and vacuum estimates versus 3 engine power
settings, and is the same as the one that
is also provided for the “orbit basics+” course series. The third one (“r noz alt”) is the same as “rocket
noz”, but with an additional calculation
block of thrust and specific impulse vs altitudes from sea level to vacuum. This is for the 3 power settings, and it creates a couple of plots. There are no inputs, it is automatic.<o:p></o:p></p><p class="MsoNormal"><b>Figure 5</b> depicts an image of the basic calculation block
that is in both “rocket noz” and “r noz alt”.
User inputs are highlighted yellow,
and significant results highlighted blue. There are two other inputs that you must deal
with, after putting the basics into the main
input block. <o:p></o:p></p><p class="MsoNormal">One of these is the design chamber pressure Pc value for
sizing the expansion ratio, and it is
not necessarily the max value. It works
with the exit plane expanded pressure Pe to size the nozzle expansion ratio
A/A*. If you are designing to a fixed
expansion, vary Pe iteratively to obtain
the desired value of A/A*. If instead
you are designing to an incipient nozzle separation situation at one of the
power settings, iteratively set Pe until
you get the desired separation backpressure Psep in that cell for that power
setting (this Psep is usually <u>just barely above</u> standard sea level
pressure). <o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal">The other is the appropriate thrust coefficient (sea level C<sub>F</sub>
or out-in-vacuum C<sub>Fvac</sub>) from the sized expansion, to use with the thrust requirement input for
sizing dimensions and flow rates. This
determines whether you size to that thrust level at sea level or in
vacuum. The rest is automatic.</p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhwahg7G8CmCaENnB02EwZ9O4Mbzzz9N8sB6fdU0K8X1WrPQkbev3AW3XvM4hLERKo7K9j37jhJSimQIkWYRatdaJperWXqqCPz8OaGqANDh6e6ZiUdt70zokyp3tYcwt5od80VDAyLLLtgnMp7jj8xatUWHFW0Hp5olRZWyYuB6PTqeB87Ju1ftedgxaXY/s1764/F5%20basic%20calc%20block.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="555" data-original-width="1764" height="126" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhwahg7G8CmCaENnB02EwZ9O4Mbzzz9N8sB6fdU0K8X1WrPQkbev3AW3XvM4hLERKo7K9j37jhJSimQIkWYRatdaJperWXqqCPz8OaGqANDh6e6ZiUdt70zokyp3tYcwt5od80VDAyLLLtgnMp7jj8xatUWHFW0Hp5olRZWyYuB6PTqeB87Ju1ftedgxaXY/w400-h126/F5%20basic%20calc%20block.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 5 – Basic Calculation Block In Both “rocket noz” and “r
noz alt” Worksheets<o:p></o:p></p><p class="MsoNormal">The altitude calculation block that is only included in the “r
noz alt” worksheet is depicted in <b>Figure 6</b>. The first 3 columns show altitude in 1000’s
of feet (kft), the ambient pressure
ratio to standard sea level pressure,
and the ambient pressure Pa in psia.
Pa is set to zero at 300 kft (about 90 km) to represent vacuum. There’s a group of columns for each of the
input power settings, in which the
vacuum thrust is corrected to thrust-at-that-altitude with F = Fvac – Pa*Ae, and the Isp computed from F and the total
flow rate at that power setting. <b><i>The
separation backpressure is also computed for that power setting.</i></b> <o:p></o:p></p><p class="MsoNormal">The user is cautioned to look at the separation backpressure
levels and compare those to the ambient pressures at altitude. Wherever the ambient pressure exceeds the
separation backpressure, the bell will
separate at that altitude and power setting.
The plots are generated without regard to separation. There will be a critical altitude below which
separation occurs, if it occurs at
all. That “trigger altitude” will be
different for each power setting. <o:p></o:p></p><p class="MsoNormal">Each power setting’s specific impulse values are summed and
then divided by the number of entries in the table. This is an approximation-only to the true
ascent-averaged specific impulse at that power setting, but it is “in the ballpark”. Doing it this way ignores the fact that the
vehicle does not spend equal times at each altitude. This is partly made up for, by there being denser points at lower
altitudes. <o:p></o:p></p><p class="MsoNormal">
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgcfpfRx0ZIjr1X1mDDgNVi0eqyWHuOIkuCod2pRycXfK3xDAqPandDuCtNUHsoeIl91m1ZE3ToPc4NwRF1SPR2OkANICJLnCIyItfmpwVsAR0wEze76EaZQsawMLlpAPjxGWqR2SEKSa1xTMXxx2jvklUfMHpf94JymEwydicE-f2HlHG1utQhefM3JSrz/s1441/F6%20alt%20calc%20block.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="889" data-original-width="1441" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgcfpfRx0ZIjr1X1mDDgNVi0eqyWHuOIkuCod2pRycXfK3xDAqPandDuCtNUHsoeIl91m1ZE3ToPc4NwRF1SPR2OkANICJLnCIyItfmpwVsAR0wEze76EaZQsawMLlpAPjxGWqR2SEKSa1xTMXxx2jvklUfMHpf94JymEwydicE-f2HlHG1utQhefM3JSrz/w400-h246/F6%20alt%20calc%20block.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 6 – The Altitude Calculation Block Only In “r noz alt”
Supporting Plots vs Altitude<o:p></o:p></p><p class="MsoNormal">If you want a copy of this spreadsheet file, please contact me. There is no user manual for it, though.
This article is probably user manual-enough.<o:p></o:p></p><p class="MsoNormal"><b><span face=""Aptos",sans-serif" style="font-size: 11pt; line-height: 107%; mso-ansi-language: EN-US; mso-ascii-theme-font: minor-latin; mso-bidi-font-family: "Times New Roman"; mso-bidi-language: AR-SA; mso-bidi-theme-font: minor-bidi; mso-fareast-font-family: Aptos; mso-fareast-language: EN-US; mso-fareast-theme-font: minor-latin; mso-hansi-theme-font: minor-latin;">Update
3-5-2024 for -- Launch to Low Earth Orbit:
Fixed-Geometry Options</span></b></p><p class="MsoNormal"></p><p class="MsoNormal">In looking at what I posted,<span style="mso-spacerun: yes;">
</span>I noticed an error and two lacks.<span style="mso-spacerun: yes;">
</span><span style="background: yellow; mso-highlight: yellow;">The error was
indicating a 2000 psia max Pc in the figures depicting the engines,<span style="mso-spacerun: yes;"> </span>when what I actually used was 2500 psia as
the max Pc.</span><span style="mso-spacerun: yes;"> </span>That has been
corrected in the versions of those figures included below.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">The two lacks were (1) not including the sizes of the
engines in the figures,<span style="mso-spacerun: yes;"> </span>and (2) not
looking at even higher vacuum design exit area ratios A/A* than the 100 that
was in the article as originally posted.<span style="mso-spacerun: yes;">
</span><span style="background: yellow; mso-highlight: yellow;">Both lacks have
been fixed with this update.</span><span style="mso-spacerun: yes;"> </span>I
added A/A* = 150 and 200 as Figures 2B and 2C,<span style="mso-spacerun: yes;">
</span>and to the comparison table.<span style="mso-spacerun: yes;">
</span>Plus,<span style="mso-spacerun: yes;"> </span>I included bell exit
diameters and estimated lengths (throat to exit) in the comparison table,<span style="mso-spacerun: yes;"> </span>now included as Figure 4 in this update.<span style="mso-spacerun: yes;"> </span>I also further annotated that comparison
table.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Bear in mind that all of these designs <u>share the same
modest-modern technology design characteristics</u>:<span style="mso-spacerun: yes;"> </span>max Pc = 2500 psia,<span style="mso-spacerun: yes;"> </span>pressure turndown ratio in throttling is
2.5:1,<span style="mso-spacerun: yes;"> </span>for 40% min pressure = 1000 psia,<span style="mso-spacerun: yes;"> </span>a 2% dumped bleed fraction,<span style="mso-spacerun: yes;"> </span>an 18-8-degree curved bell profile,<span style="mso-spacerun: yes;"> </span>and a nozzle throat discharge efficiency of
99.5%.<span style="mso-spacerun: yes;"> </span>They are all “paper” oxygen-hydrogen
(LOX-LH2) engines.<span style="mso-spacerun: yes;"> </span>They do <u>not</u>
push the state-of-the-art very hard!<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal"><b>Figure 1 below</b> is the same as Figure <a name="_Hlk160541979">1 in the original article above,<span style="mso-spacerun: yes;"> </span>except that I corrected the error reporting
max Pc,<span style="mso-spacerun: yes;"> </span>and I added bell length.</a><span style="mso-spacerun: yes;"> </span>This <b>sea level design</b> sizes its
expansion between 2500 psia Pc,<span style="mso-spacerun: yes;"> </span>and
14.696 psia Pe = Pa at sea level.<span style="mso-spacerun: yes;"> </span>The
dimensions and flow rates size for 500,000 lb of thrust delivered at sea
level.<span style="mso-spacerun: yes;"> </span>It is neither over-expanded nor
under-expanded.<span style="mso-spacerun: yes;"> </span>It is “perfectly
expanded”,<span style="mso-spacerun: yes;"> </span><u>as any real “sea level
design” actually is</u>. <o:p></o:p></p>
<p class="MsoNormal"><b>Figure 2 below</b> is the same as Figure 2 in the
original article above,<span style="mso-spacerun: yes;"> </span>except that I
corrected the error reporting max Pc,<span style="mso-spacerun: yes;">
</span>and I added bell length.<span style="mso-spacerun: yes;"> </span>This is
a “vacuum design” sized to an arbitrary A/A* = 100.<span style="mso-spacerun: yes;"> </span>The dimensions and flow rates size for an
imposed vacuum thrust requirement of 500,000 lb.<span style="mso-spacerun: yes;"> </span>Under- or over-expandedness at sea level is
irrelevant,<span style="mso-spacerun: yes;"> </span>since this design is
separated at sea level,<span style="mso-spacerun: yes;"> </span>at any throttle
setting.<span style="mso-spacerun: yes;"> </span>It cannot be used for ascent,<span style="mso-spacerun: yes;"> </span>except as an upper stage engine used only
essentially exoatmospherically.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal"><b>Figure 2B below</b> is the same as the corrected Figure 2
reported here in every way,<span style="mso-spacerun: yes;"> </span>excepting
only that the design A/A* = 150,<span style="mso-spacerun: yes;"> </span>instead
of Figure 2’s 100.<span style="mso-spacerun: yes;"> </span><b>Figure 2C below</b>
is the exact same thing yet again,<span style="mso-spacerun: yes;">
</span>except that the design expansion ratio is A/A* = 200.<span style="mso-spacerun: yes;"> </span>These were added to investigate the effects
of higher expansion ratio upon the performance and dimensions of vacuum
engines. <o:p></o:p></p>
<p class="MsoNormal"><b>Figure 3 below</b> is the same as Figure 3 in the
original article above,<span style="mso-spacerun: yes;"> </span>except that I
corrected the error reporting max Pc,<span style="mso-spacerun: yes;">
</span>and I added bell length.<span style="mso-spacerun: yes;"> </span>This one
is my “compromise design”,<span style="mso-spacerun: yes;"> </span>in which I
trade an over-expanded loss of Isp performance right at sea level,<span style="mso-spacerun: yes;"> </span>for <u>both</u> higher vacuum Isp,<span style="mso-spacerun: yes;"> </span><u>and</u> a higher ascent-averaged Isp,<span style="mso-spacerun: yes;"> </span>than a traditional sea level design. <span style="mso-spacerun: yes;"> </span>It is a slightly-larger engine,<span style="mso-spacerun: yes;"> </span>though. <o:p></o:p></p>
<p class="MsoNormal"><b>Figure 4 below</b> is the Table in the original article
above,<span style="mso-spacerun: yes;"> </span>enlarged, <span style="mso-spacerun: yes;"> </span>and annotated for the important overall design
considerations that impact vehicle design the most.<span style="mso-spacerun: yes;"> </span>The basic data in the original table are
unchanged,<span style="mso-spacerun: yes;"> </span>just rows added representing
the two vacuum designs with the larger expansion ratios.<span style="mso-spacerun: yes;"> </span>Columns have been added on the right with
values for A/A*,<span style="mso-spacerun: yes;"> </span>exit diameter De,<span style="mso-spacerun: yes;"> </span>and an estimate of bell length from throat to
exit plane.<span style="mso-spacerun: yes;"> </span><span style="mso-spacerun: yes;"> </span>Annotations have been added as to what is most
important for designing realistic vehicles. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjZsplJeL7KRTq8d7yBtwz5xLiTFvRgn3WBat5KjlbQobTRwIAa1ZXGxGnejdt-_em8qDw-qPiOreGIGcnFpv7fDHqU3AGkYS0cx5R3DCi1BCRxfV5HdcjSCFgOzaR1LhDU2Xbqv5oXQRWqgO6Gm3Tu-zElOql7WUBPjy-bZr22uYD5fnUibcqsFCDOccw0/s975/F1%20sea%20level.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="613" data-original-width="975" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjZsplJeL7KRTq8d7yBtwz5xLiTFvRgn3WBat5KjlbQobTRwIAa1ZXGxGnejdt-_em8qDw-qPiOreGIGcnFpv7fDHqU3AGkYS0cx5R3DCi1BCRxfV5HdcjSCFgOzaR1LhDU2Xbqv5oXQRWqgO6Gm3Tu-zElOql7WUBPjy-bZr22uYD5fnUibcqsFCDOccw0/w400-h251/F1%20sea%20level.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 1 – The Baseline Sea Level Design, Perfectly-Expanded At Sea Level<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhkZ8tQGt99QF-00yBLfNqK_3u8UtPjkqimWuU_MQD-HEey7P229QncF04Rvd3AOk_rcvBfFTaep7nhgz4f_M8UgSFTtiwz5G2j72PUdSwmH4wkO4hLYBJRHSxvdEEzPgM9lCnbuXv3VnG6EMp4691_rFgHZ1QP69VeX3lvoKUYQkzgSdYZMt8NjaBrIq5C/s973/F2%20vacuum.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="973" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhkZ8tQGt99QF-00yBLfNqK_3u8UtPjkqimWuU_MQD-HEey7P229QncF04Rvd3AOk_rcvBfFTaep7nhgz4f_M8UgSFTtiwz5G2j72PUdSwmH4wkO4hLYBJRHSxvdEEzPgM9lCnbuXv3VnG6EMp4691_rFgHZ1QP69VeX3lvoKUYQkzgSdYZMt8NjaBrIq5C/w400-h251/F2%20vacuum.png" width="400" /></a></div><p></p><p class="MsoNormal"><a name="_Hlk160544010">Figure 2 – The Baseline Vacuum
Design, Unusable For Ascent, Sized At A/A* = 100<o:p></o:p></a></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg7-7goE_aSAtcOpJRgIACAuv9PV1x-Ptnk6aPR8YFFseWwu8gwTaNn3dMEnFDjvMuqk39fNLLYgKFDYmVEt4ifFI5AgnWHDlzRb036ZzuG2H1w_ofoGiOyyfC0Zi1CvzmuS13KFq9x_ThC6CFSQ3PQgf_qIbOix3vaXIi7LNDo8e_uTCVuZ1D0i9u5eY1d/s974/F2B%20vacuum%20150.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="609" data-original-width="974" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg7-7goE_aSAtcOpJRgIACAuv9PV1x-Ptnk6aPR8YFFseWwu8gwTaNn3dMEnFDjvMuqk39fNLLYgKFDYmVEt4ifFI5AgnWHDlzRb036ZzuG2H1w_ofoGiOyyfC0Zi1CvzmuS13KFq9x_ThC6CFSQ3PQgf_qIbOix3vaXIi7LNDo8e_uTCVuZ1D0i9u5eY1d/w400-h250/F2B%20vacuum%20150.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2B – A Vacuum Design,
Unusable For Ascent, Sized At
A/A* = 150<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgjq1NGuO3pR0nM8Oi2fBbgc7AdgYC9YHjB88x_AFHNU2HlZmFm8BhDk0S-u6P8Rwl7E8bX4MIAMD70RnU4-XXwdbragrMqGCxhc2Iwf4T7EY7sk3wTC1HStoROoHLE1cFFtV8_zQz_uPTDKbP4T-ih4abD0JFFaQ4WvFypWyr0Rl16min-0vb-cHxwg6mH/s976/F2C%20vacuum%20200.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="611" data-original-width="976" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgjq1NGuO3pR0nM8Oi2fBbgc7AdgYC9YHjB88x_AFHNU2HlZmFm8BhDk0S-u6P8Rwl7E8bX4MIAMD70RnU4-XXwdbragrMqGCxhc2Iwf4T7EY7sk3wTC1HStoROoHLE1cFFtV8_zQz_uPTDKbP4T-ih4abD0JFFaQ4WvFypWyr0Rl16min-0vb-cHxwg6mH/w400-h250/F2C%20vacuum%20200.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2C – A Vacuum Design,
Unusable For Ascent, Sized At
A/A* = 200<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhJmgt7VJNzbQFee378ajevOxrIFFdYAFpv24ghpUdPkMizPD-Yoyu5ZRVjmgvJyqpJXdmrqGPnfEbwd2rDrIGMX4nCnOe1McmQwS4IxMoxqfqzCnI_bp5zwqIlY0YLzU63sU5oJwaYz87RGMclW81WmzxngtNRANXTeN-61dznJ1otUgfWeXgRoR8KWGuX/s977/F3%20underexpanded.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="613" data-original-width="977" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhJmgt7VJNzbQFee378ajevOxrIFFdYAFpv24ghpUdPkMizPD-Yoyu5ZRVjmgvJyqpJXdmrqGPnfEbwd2rDrIGMX4nCnOe1McmQwS4IxMoxqfqzCnI_bp5zwqIlY0YLzU63sU5oJwaYz87RGMclW81WmzxngtNRANXTeN-61dznJ1otUgfWeXgRoR8KWGuX/w400-h251/F3%20underexpanded.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 3 – The “Compromise” Design, Usable For Ascent, At a Better Ascent-Averaged Isp <o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg5iUCIUOGKs6EsCNF0DFTpybngxQAJQ6AZzPNrP2lHDJXNhTKaTB-ONidDrGGX3T7etQVsX7HyfKHZfNyb1D3ESZhT0RmS8aq9oA7gMky8VDXTCU5gotp7ccZZDoYh_e-cr9QFx2eMooNIRtrXRVGwnIPxIS368awbYv3M2_lcnV6By7CEL9h-SYGRRLrA/s1162/table%202.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="569" data-original-width="1162" height="196" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg5iUCIUOGKs6EsCNF0DFTpybngxQAJQ6AZzPNrP2lHDJXNhTKaTB-ONidDrGGX3T7etQVsX7HyfKHZfNyb1D3ESZhT0RmS8aq9oA7gMky8VDXTCU5gotp7ccZZDoYh_e-cr9QFx2eMooNIRtrXRVGwnIPxIS368awbYv3M2_lcnV6By7CEL9h-SYGRRLrA/w400-h196/table%202.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 4 – The Revised Comparison Table, Annotated For Important Overall Design
Considerations<o:p></o:p></p><p class="MsoNormal">The “compromise” engine is not quite as small as a true sea
level design, but it outperforms the sea
level design for ascents overall, and
this would apply whether single or 2-stage. <o:p></o:p></p><p class="MsoNormal">It is not as large as the three vacuum engines, although it is fairly close in size to the
A/A* = 100 design, and in its vacuum
performance. <o:p></o:p></p><p class="MsoNormal">It is much smaller than the A/A* = 150 and = 200
designs, and yet still does not fall all
that far short, in terms of vacuum
performance (under 5% shortfall). A
factor of 1.050 on the velocity ratio dV/Vex is only a factor of about 1.051 applied
between the corresponding mass ratios needed for vehicle or stage design. <o:p></o:p></p><p class="MsoNormal">The compromise design is like the other fixed-geometry
designs: it will have a very high engine
thrust/weight ratio T/We ~ 50-100. <u>That
<b><i>cannot</i></b> be true of anything that is variable geometry</u>!<o:p></o:p></p><p class="MsoNormal"><b>Conclusions<o:p></o:p></b></p><p class="MsoNormal"><b>#1.</b> If you are doing a 2-stage launch vehicle design
(first stage expendable or reusable),
then use the “compromise design” approach to size your first stage
engines. <u>They will outperform a
conventional perfectly-expanded sea level design</u>! The downside is very limited throttle-down
capability at sea level. <o:p></o:p></p><p class="MsoNormal">Use a “vacuum design” at the largest dimensions you can
tolerate, for your expendable second
stage engines, since the stage point
will always be essentially exoatmospheric,
and very near to horizontal flight.
<o:p></o:p></p><p class="MsoNormal">If your second stage is reusable, it is unclear <u>whether </u>you need a mix
of sea level and vacuum engines, <u>or</u>
you could just use engines designed with the “compromise design” approach. The deciding factor will be the minimum
thrust per engine to actually land. <o:p></o:p></p><p class="MsoNormal">Vacuum designs cannot be used for this, because they are always flow-separated near
sea level. The “compromise” approach may
be separated, if you have to throttle
below its expansion design point. <u>Only
the true sea level designs will be unseparated at both sea level and min
throttle setting</u>. <o:p></o:p></p><p class="MsoNormal"><b>#2.</b> If you are doing some sort of single-stage to
orbit design, use the “compromise
design” approach in preference to the traditional perfectly-expanded sea level
engine approach. This will get you a
higher ascent-averaged Isp than the traditional sea level designs can
achieve. <o:p></o:p></p><p class="MsoNormal">If this is to be a reusable design that lands vertically, the problem is at landing, when the weight is lowest. It is launch that sets the summed thrust of
all engines, with due allowance for an
engine or two nonfunctional. At
landing, you need to be able to shut
down enough of them so that the remaining engine or engines can be throttled
down enough to land, without suffering
separation. <o:p></o:p></p><p class="MsoNormal">What made the “compromise design” do better in ascent than a
sea level design, was sizing its
expansion in the vicinity of 85% of max Pc.
That is a very small throttle range for vertical landing. You may instead have to include some sea
level engines for landing, that can
throttle deeply. <o:p></o:p></p><p class="MsoNormal">The way to avoid this issue is to land horizontally, so that thrust at touchdown is not
needed, excepting perhaps a “go-around”
capability. The downside to this is <u>far
higher stage inert fraction</u>,
inherent because of the required lifting shape, whether winged, or as a lifting body. Exclusive of tankage and engine masses, about the min credible airframe mass fraction
will be near 10% or so. The engines and
the propellant tanks add directly to that!<o:p></o:p></p><p class="MsoNormal"><b>#3.</b> Why make this more difficult with heavier
variable-geometry engines, with a longer
list of failure modes? Why incur the low
vacuum performance of free-expansion designs?
Neither option makes any practical sense.<o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal">------<o:p></o:p></p><p class="MsoNormal">Just in case you do not understand why free-expansion
designs have lousy vacuum performance,
see <b>Figure 5</b>. The nozzle
kinetic energy efficiency reflects only the integrated average of the cosine
factors for exiting streamlines that are not aligned with the thrust axis. Such is measured after the last point of
contact, not before! Thrust
is measured just before the last point of contact, not after!
This applies only to the momentum term m*Ve of thrust. The as-expanded Pe*Ae term is usually smaller
by far, and does not get ratioed by the
kinetic energy efficiency. <o:p></o:p></p><p class="MsoNormal">The usual nozzle average of cosine factors is literally the
average of 1 on the centerline, and
whatever the outer-edge streamline cosine factor is. That simple arithmetic average may not reflect
the true integrated average. However, it is still somewhere in the ballpark! And because of Prandtl-Meyer expansion
effects at the outer edge of the exiting flow,
at high altitudes, that edge cosine
factor is very near-zero, or even
slightly negative! <span style="background: yellow; mso-highlight: yellow;">Thus the nozzle kinetic energy
efficiency is catastrophically low! Almost no matter how you figure it!</span><o:p></o:p></p><p class="MsoNormal">And <u>THAT </u>is exactly why free-expansion nozzle designs
of <u>ANY</u> kind, are truly <u>LOUSY</u>
vacuum engine designs! They work better
than conventional bells up to the lower stratosphere, but they inherently degrade very quickly into
uselessness, much above the lower
stratosphere. <o:p></o:p></p><p class="MsoNormal">I know that many perceive it differently, but you have been lied to, for marketing purposes, in a corporate welfare system. Look instead at the actual engineering numbers. And you will need to understand what
Prandtl-Meyer expansion is, in
compressible flow, to fully make sense
of this. <o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal"><o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjhZktIxmA7FmaxDDpmL_vDVAMlCUCBUkxPRBUjOM2sQTfF9sNnkT0YCMOw4-r92XaTTfLSBbTJ_VxNoUKJ1j74618IlwyZzRHOfrKoz36lzbWPoGax5cql8yQj1qsk0E5kICcGMTgfJHEaAKVyaR7ku0CKgw5ZKa337G3RlI4zTwlLugUMx5T9hFgi53Ua/s1077/F5%20free%20expansion.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="731" data-original-width="1077" height="271" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjhZktIxmA7FmaxDDpmL_vDVAMlCUCBUkxPRBUjOM2sQTfF9sNnkT0YCMOw4-r92XaTTfLSBbTJ_VxNoUKJ1j74618IlwyZzRHOfrKoz36lzbWPoGax5cql8yQj1qsk0E5kICcGMTgfJHEaAKVyaR7ku0CKgw5ZKa337G3RlI4zTwlLugUMx5T9hFgi53Ua/w400-h271/F5%20free%20expansion.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 5 – Why Free-Expansion Nozzle Designs Always Have
Lousy Vacuum Performance<o:p></o:p></p><p class="MsoNormal"><br /></p><b><span face=""Aptos",sans-serif" style="font-size: 11pt; line-height: 107%; mso-ansi-language: EN-US; mso-ascii-theme-font: minor-latin; mso-bidi-font-family: "Times New Roman"; mso-bidi-language: AR-SA; mso-bidi-theme-font: minor-bidi; mso-fareast-font-family: Aptos; mso-fareast-language: EN-US; mso-fareast-theme-font: minor-latin; mso-hansi-theme-font: minor-latin;"></span></b><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com2tag:blogger.com,1999:blog-2675974463524895416.post-35135247504215954882024-03-03T10:27:00.003-06:002024-03-07T15:31:54.010-06:00Launch to Low Earth Orbit: 1 or 2 Stages?<p class="MsoNormal">Although I have examined this question before, I wanted to look at it again, because there is still enthusiasm for the
single-stage notion using chemical propulsion.
The problem with that is achieving a very high specific impulse (Isp)
across a broad range of altitudes with the stage engines. These <u>must have adequate thrust at sea
level</u>, but also average a high Isp
all across the ascent. Those
requirements are in conflict because of fundamental physics.<o:p></o:p></p>
<p class="MsoNormal">The two-stage notion does not face that quandary as
directly. While the first stage engines
show the reduced Isp typical of a sea level design, which does not improve much at all going to
high altitudes, that penalty is
compensated by the first stage shouldering <u>only a minority fraction</u> of
the total delta-vee (dV) requirement to low Earth orbit (LEO). The second stage can be a “vacuum”
design, featuring much higher Isp, with much-relaxed thrust requirements. <o:p></o:p></p>
<p class="MsoNormal">Fixed-geometry rocket engines provide the shortest list of
possible failure modes, compared to
variable -geometry designs that compensate by deployable expansion bell
extensions. Fixed-geometry rockets also
show much higher performance out in actual vacuum than any of the
free-expansion designs, because of the
very high streamline divergence the free-expansion designs inherently suffer
when out in actual vacuum. (They work
“best” in the lower stratosphere.)<o:p></o:p></p>
<p class="MsoNormal">Accordingly, what I
looked at here were entirely fixed-geometry rocket engines. For the two-stage notion, the first stage engines were sea level
designs, and the second stage engines
were “vacuum” designs, although, strictly speaking, there is no such thing as a “vacuum”
design, there are only practical design
constraints on how big the expansion can actually be (it has to fit behind the
stage). <o:p></o:p></p>
<p class="MsoNormal">The sea level designs size the expansion ratio to be
perfectly-expanded at sea level for no pressure term penalty, and its dimensions also size there, <u>to meet a sea level max thrust requirement</u>.
This is because the vehicle is the heaviest at ignition, and yet adequate net acceleration upward
against gravity (around half a gee net) must be obtained! In addition,
the sea level engines were presumed to use kerosene-oxygen
propellants, in order to minimize first
stage tankage volume and frontal area,
so that drag losses are minimized.
Not to do so makes the needed mass ratio even larger. <o:p></o:p></p>
<p class="MsoNormal">The vacuum designs for second stages (and for the
single-stage design) were presumed designed for expansion just short of
backpressure-induced flow separation at sea level, at a suitable part-throttle
condition: some 85% of max chamber
pressure. In that way, the actual flight engines can be tested
open-air nozzle at sea level, at
85%-and-above chamber pressure,
drastically reducing development test costs! Similarly, in flight, thrust can be reduced to 85% chamber pressure
levels from sea level on up, without
risking flow separation in the expansion bell.
<o:p></o:p></p>
<p class="MsoNormal">For the two-stage design,
a vacuum thrust requirement can be used to set dimensions. This second stage was presumed to use
oxygen-hydrogen propellants, since the
smaller stage volume is compatible with the same or smaller frontal area, despite the low density of liquid
hydrogen. For the single-stage
design, <u>a sea level thrust
requirement must be imposed</u>. The
single stage design needs a higher-energy propellant combination, but also suffers greatly from the enormous
tankage volumes and frontal area of a hydrogen design. So, a
compromise was used:
methane-oxygen. <o:p></o:p></p>
<p class="MsoNormal"><o:p></o:p></p><p>The launch trajectory was presumed to be a thrusting gravity
turn affected by atmospheric drag, to
LEO at low inclination eastward, as
shown in <b>Figure 1</b>. For the <u>all-expendable
designs presumed here</u>, staging would
be somewhere near 50 km altitude and about 2 km/s achieved speed. Circular orbit speed near 300 km altitude is
about 7.7 km/s achieved. Assuming 5%
each for gravity and drag losses, the
mass ratio-effective dV is about 8.5 km/s.
The loss to be overcome is thus about 0.8 km/s, all assigned to the first stage of a
two-stage vehicle as a decent approximation,
and all borne by the single-stage vehicle.</p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEipWEgBfU5k4vVZlPWz_Jpdr6w4lJysM1Ji7VXC2OOC-j8vap3IC4h0fi92VGlorktocS8UccIi-4gGTGLQNfBtSONIvJ13_BwfMxi11K4z2toCSalXQqeBKwrUB1o9OO4ctWblEd6dWnhfaAusptE_E6rzodK8d9_tFajSCnrC-RuTGtZY8juH0WpHTSfg/s977/F1%20launch%20reqmnts.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEipWEgBfU5k4vVZlPWz_Jpdr6w4lJysM1Ji7VXC2OOC-j8vap3IC4h0fi92VGlorktocS8UccIi-4gGTGLQNfBtSONIvJ13_BwfMxi11K4z2toCSalXQqeBKwrUB1o9OO4ctWblEd6dWnhfaAusptE_E6rzodK8d9_tFajSCnrC-RuTGtZY8juH0WpHTSfg/w400-h250/F1%20launch%20reqmnts.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 1 – Launch Requirements<o:p></o:p></p><p class="MsoNormal">I did not actually size engines for the kerosene-oxygen and
hydrogen-oxygen engines of the two-stage design, because I have done this before, and my results match general industry
experiences. These represent only
modestly state-of-the art designs: 330 s
Isp for the kerosene-oxygen, and 450 s
Isp for the hydrogen-oxygen. These would
be for chamber pressures in the 2000-3000 psia range, and maybe 2% bleed.<o:p></o:p></p><p class="MsoNormal">I presumed a very state-of-the-art methane-oxygen engine of
full flow cycle so that bleed was zero,
with a very high max chamber pressure of 4000 psia and a rather-demanding
pressure turndown ratio (P-TDR) of 3. I
also presumed I would size its expansion from 85% chamber pressure down to 3.3
psia, with the separation-inducing
backpressure set at 14.70 psia. For
initial rough-sizing purposes, I simply presumed
it would average 370 s Isp across its full ascent. <o:p></o:p></p><p class="MsoNormal">For vehicle rough-out sizing, I presumed a 5% inert fraction (f<sub>inert</sub>)
for all stages, as loaded with
payload. The payload was presumed to be
a dead-head 100 metric tons,
streamline-shaped, and mounted
out in the open, atop the launch
vehicle. The ratio of dV to effective
exhaust velocity (Vex) determines the stage mass ratio (MR). The propellant mass fraction (f<sub>prop</sub>)
of the loaded stage is then 1 – 1/MR.
And the payload fraction (f<sub>pay</sub>) is thus 1 – f<sub>prop</sub>
– f<sub>inert</sub>. For the two-stage
launch vehicle, the first stage
“payload” is the fully loaded and fueled second stage mass. <o:p></o:p></p><p class="MsoNormal">I used a very
simply laid-out spreadsheet to calculate these numbers for the two
designs, using the presumed Isp values
and the relationship Vex = gc*Isp/1000,
to get Vex in km/s to match the dV values. Those initial results are shown in <b>Figure
2</b>. Note that the one-stage design
has about half the overall payload fraction and twice the launch mass of the
two-stage design! I used thrust-to-weight
(T/W) ratios of 1.5 at liftoff for good ascent kinematics, and a T/W just over 1 for the
exo-atmospheric, nearly-horizontal
portion of flight near the end of the ascent.
These sized some stage thrust requirements for me. I used only half-a-gee for the second stage
of the two-stage vehicle.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgGtF3JTyHlqXzWjX79mr3yTiwlqItEelQz3uB6jpMik_u_MQEuxxKrO2Vu0oq7TdF1nGGFfI7iOZvpYDfvssXYjfEzaOWxJDhX3kG8_faFLQ2WoNK8pah22_aHvPYb40ZxVX4bib5NX7EOWb9gLIn1ZdZ02VLx1YTxeLiSpCEbstUcjOwabI_phkG4XTGR/s977/F2%20veh%20rough.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgGtF3JTyHlqXzWjX79mr3yTiwlqItEelQz3uB6jpMik_u_MQEuxxKrO2Vu0oq7TdF1nGGFfI7iOZvpYDfvssXYjfEzaOWxJDhX3kG8_faFLQ2WoNK8pah22_aHvPYb40ZxVX4bib5NX7EOWb9gLIn1ZdZ02VLx1YTxeLiSpCEbstUcjOwabI_phkG4XTGR/w400-h250/F2%20veh%20rough.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2 – First Vehicle Rough-Out<o:p></o:p></p><p class="MsoNormal"> <b><i>Revisiting
the Rough-Out<o:p></o:p></i></b></p><p class="MsoNormal">I really had no questions regarding the feasibility of the
presumed engine Isp levels for the two-stage design. There was concern about the Isp = 370 s
presumption for the one-stage design.
Accordingly, I actually ran some
engine sizing and performance estimates,
using a convenient spreadsheet tool.
The ascent-averaged Isp fell closer to 360 s than the initially-presumed
370 s. This is illustrated in <b>Figure
3</b> below. That includes a sketch and
notations, plus some copied sizing and
point performance data from the spreadsheet.
The predicted performance vs altitude plots from that same spreadsheet are
given in <b>Figure 4</b> below. <o:p></o:p></p><p class="MsoNormal">To find the ascent-averaged Isp from the calculation block
in the spreadsheet, which is performance
vs altitude, I simply summed the 100% Isp
values over the ascent, and divided that
by the number of entries in the table.
This is not the “right” average value,
because the vehicle does not spend equal amounts of time at each
altitude, but it is somewhere in the
ballpark. The Isp out in vacuum is
pretty near the initial presumption of 370 s Isp, but the low altitude values are much
lower, and the vehicle does spend a lot
of time there, since it is still moving
slowly at low altitude. <o:p></o:p></p><p class="MsoNormal">Therefore, I reran
the vehicle size-out and thrust requirements for that one-stage vehicle, with the nominally-lower presumed average Isp
= 360 s. That revised vehicle rough-out
is depicted in <b>Figure 5</b> below,
which is just Figure 2 edited in some places. The edits are in red text. The effect of the small Isp change is more
dramatic on the one-stage vehicle than it would be for either stage of the
2-stage vehicle. This is precisely
because it is only one stage, and the
payload is a fixed number.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgBGLrH20GHrC0AO92uUHbV_HiaJvJavBsv3PYKjmq0krG4ZjqNN_2Syg-JIFpRSH1zj5y5yFFjBF7L2FBQ8zAHVK5ZXplXYanD9i2WsvzzmpEk4urqEF-n3qic-bLK5XyrUGAsb1B3ATnQxlqEZAvNM_gPp0Ft9K0xqWry5pKCRGZceyLwnwOU9OgplemF/s977/F3%20eng%20sz.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgBGLrH20GHrC0AO92uUHbV_HiaJvJavBsv3PYKjmq0krG4ZjqNN_2Syg-JIFpRSH1zj5y5yFFjBF7L2FBQ8zAHVK5ZXplXYanD9i2WsvzzmpEk4urqEF-n3qic-bLK5XyrUGAsb1B3ATnQxlqEZAvNM_gPp0Ft9K0xqWry5pKCRGZceyLwnwOU9OgplemF/w400-h250/F3%20eng%20sz.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 3 – LOX-LCH4 Engine Sizing (Re-scalable With Thrust
Rating)<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhyLCRMDjE3GG95VqPD5S-4A5nY3nfj0iGTqZST5kyOvWeIYe2JBlSsbzPuBW5THR6GJxeP38xk2NqbYackfMbL0hnxYe1cJU5QqBjdrt665yCxc7xuRxp8deDzqyTqAkWP__og5zuXYKkVt_nSeK6nzkxS_ZEhyphenhyphen302BAeDRCKA9IIVnI-3UGTaMNpCQy2p/s1132/F4%20perf.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="651" data-original-width="1132" height="230" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhyLCRMDjE3GG95VqPD5S-4A5nY3nfj0iGTqZST5kyOvWeIYe2JBlSsbzPuBW5THR6GJxeP38xk2NqbYackfMbL0hnxYe1cJU5QqBjdrt665yCxc7xuRxp8deDzqyTqAkWP__og5zuXYKkVt_nSeK6nzkxS_ZEhyphenhyphen302BAeDRCKA9IIVnI-3UGTaMNpCQy2p/w400-h230/F4%20perf.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 4 – Predicted LOX-LCH4 Performance Fell Short<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEheBgd9WyTIQPSHmK9v4aCZGsorq4gJb-vPo3yjzTcMs8nEOMA15hcuGV0SFMs9tsTWkRGFhebDK0ltjhD5H0oizlMBAn8x4UHi27GJOZCCImQD9P1z9QlHYqoaUTOJREZ1ouTWx-bY_sKoL5RJgL9QpWt4KYd5F6uuqlGFjW0iha0WS1WQbN7JPJpshQoU/s977/F5%20veh%20rough%202.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEheBgd9WyTIQPSHmK9v4aCZGsorq4gJb-vPo3yjzTcMs8nEOMA15hcuGV0SFMs9tsTWkRGFhebDK0ltjhD5H0oizlMBAn8x4UHi27GJOZCCImQD9P1z9QlHYqoaUTOJREZ1ouTWx-bY_sKoL5RJgL9QpWt4KYd5F6uuqlGFjW0iha0WS1WQbN7JPJpshQoU/w400-h250/F5%20veh%20rough%202.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 5 – Revised Vehicle Rough-Out Reflects Revised Isp
For the Single-Stage Engines<o:p></o:p></p><p class="MsoNormal">This second version of the vehicle rough-out is more
reliable, after revising the one-stage
average Isp value. The two-stage vehicle
is probably “pretty close” as it is,
especially since the second stage Isp is likely a slight
underestimate, which would offset any
over-estimate of the first stage Isp. <o:p></o:p></p><p class="MsoNormal">For the two-stage vehicle,
we are probably looking at 8 or 9 engines of some 220 metric tons-force
thrust each, in the first stage. The second stage needs very little pathwise
acceleration capability, and most of
that at ignition where it is heaviest,
so the same 220 metric tons-force of thrust would work, although for redundancy, I would recommend two engines of 110 metric
tons-force thrust each. That way, it still flies adequately even if one engine
quits.<o:p></o:p></p><p class="MsoNormal">For the one-stage vehicle,
the same engines burn all the way through the ascent, only shutting down those that are not needed
as weight decreases. This is a
compromise between too many engines and too much thrust late in the
ascent. What the figure shows is that
15-16 engines of around 250 metric-tons-force each, will lift off well, with only one of those still burning very late
in the ascent. <o:p></o:p></p><p class="MsoNormal">Overall, the message
is clear: to do this one-stage cuts the
achievable payload fraction in half or less,
while increasing the liftoff mass by a factor a bit over 2, all for placing the same payload in eastward, low-inclination LEO. Lower payload fraction and higher ignition
mass increase cost!<o:p></o:p></p><p class="MsoNormal"><span style="background: yellow; mso-highlight: yellow;">The
two-stage vehicle does better, because
its two stages address the wildly-different requirements of ascent out of the
atmosphere and exo-atmospheric acceleration to orbit speed, with two entirely-different engine designs
and propellant combinations! The one-stage
design lacks that advantage, and must
push its engines to the very outer limits of the state-of-the-art.</span> <o:p></o:p></p><p class="MsoNormal" style="text-indent: 0.5in;"><b><i>Extending to Reusable
Vehicles<o:p></o:p></i></b></p><p class="MsoNormal">To do this reusably just makes the vehicles somewhat larger. For the two stage vehicle, the first stage gets larger in order to have
the extra propellant required to recover it and land it. Up to this date, there have been no demonstrations of any
recovery of second stages at all. This
is the partial recovery path taken by SpaceX with its Falcon-9 and Falcon-Heavy
vehicles. <o:p></o:p></p><p class="MsoNormal">The inert fraction of any recoverable second stage would be much
larger than the 5% presumed here,
because it must be not just a stage,
but also a survivable orbital re-entry vehicle. It might as well carry the payload
internally, which likely increases its
inert fraction even more. That path is
the one chosen by SpaceX with its Starship/Superheavy orbital transport design.
<o:p></o:p></p><p class="MsoNormal">As for making the one-stage vehicle reusable, with only 4% payload fraction, it could only have an inert fraction of
9%, even if it carried no payload at all! To make the stage also an entry vehicle, and to carry the payload internally, would seem to push well past the bounds of any
reasonable assumptions at all, with
chemical Isp. This is the path attempted
without any success by the X-33 “Venture Star” project, and it used hydrogen-oxygen, the best chemical combination available! <o:p></o:p></p><p class="MsoNormal"> <b><i>Summary
Remarks<o:p></o:p></i></b></p><p class="MsoNormal">Because launch price is sensitive to payload fraction and
ignition mass, I cannot recommend the
single-stage-to-orbit approach with any conceivable chemical propulsion, even in expendable vehicles. The numbers are just not there, regardless of what kind of “trick” engines
one proposes, because such always have
performance shortfalls somewhere across the ascent. Two-stage to orbit, using two different propellant combinations
in the two stages, is likely the
best, but SpaceX has already shown
rather good results with the same propellant combinations in both stages. <o:p></o:p></p><p class="MsoNormal">To add reusability,
the best approach is still two-stage,
with either (1) an expendable second stage and payload riding atop
it, or (2) a second stage that is also
its own entry vehicle, with payload
riding inside. The first is still more
mature than the second, at the time of
this writing. <o:p></o:p></p><p class="MsoNormal">Switching to all-hydrogen instead of the denser methane is
not the solution to the single-stage problem,
because the far-larger tankage volume and frontal area will increase the
drag loss, raising the dV penalty, and thus make mass ratio-effective dV requirement
still higher. Such acts to offset the
effects of the higher Isp of the hydrogen,
which still has to be ascent-averaged.<o:p></o:p></p><p class="MsoNormal">If you really want to do single-stage to orbit, the most fruitful thing to do would be
developing into maturity a nuclear thermal engine of significantly-higher Isp
and substantially-higher engine thrust/weight than the NERVA design that was
ready to flight test, when it was
cancelled in 1974. Such an option is
very likely some sort of gas core design.
One needs at least about Isp = 1000 s or so, to make fully-reusable stages that are their
own entry vehicle, and can contain the
payloads internally. The vehicle inert
fractions will fall in the 20-30% range,
unless high engine weight drives it even higher. The vehicle launches vertically, and could land horizontally. If clean,
dV ~ 8.5 km/s.<o:p></o:p></p><p class="MsoNormal">At 1000 s: Vex =
9.80667 km/s, MR = 2.3792, f<sub>prop</sub> = 0.5797, guess f<sub>inert</sub> = .25, f<sub>pay</sub> = 0.1703. For W<sub>pay</sub> = 100 metric tons, W<sub>ign</sub> = 587 m.ton, W<sub>inert</sub> = 147 m.ton, and W<sub>prop</sub> = 340 m.ton. At liftoff T/W = 1.5, the required liftoff thrust is 881
m.ton-force. Burnout is about 247 m.ton, for about 3.57 gees at liftoff thrust. The
thing is likely winged, or a lifting
body shape, to land on a runway or dry
lake bed.<o:p></o:p></p><p class="MsoNormal"> <b><i>Follow-Up
on the Nuclear Single-Stage Notion<o:p></o:p></i></b></p><p class="MsoNormal">I created another spreadsheet worksheet to evaluate the
possibility of a nuclear thermal one-stage design. I made the inert fraction iterative, with an R-value to estimate LH2 tankage
inerts, and an engine thrust/weight
ratio to estimate engine inerts based on liftoff thrust required. <o:p></o:p></p><p class="MsoNormal">This crude analysis includes nothing for on-orbit
maneuvering, or deorbit, which would probably be storable propellants! I made the inerts analysis iterative so that
the overall inert fraction input would give a realistic airframe inert
fraction, that does not include the
engine or the tankage. <o:p></o:p></p><p class="MsoNormal">This one is a lifting body,
with an engine not all that far improved over NERVA, and it would land dead-stick like the
shuttle, probably on a dry lakebed, or a very long runway indeed. It would likely touch down at around 200
mph. <o:p></o:p></p><p class="MsoNormal">This one had the highest payload fraction I have seen
yet, and would likely be fairly cheap to
operate, as long as it proves tolerable
to return the idled nuclear engine back to Earth. (That is a really big “if”!) See the spreadsheet image in <b>Figure 6</b>, and a sketch of the vehicle concept in <b>Figure
7 below.</b> <o:p></o:p></p><p class="MsoNormal">I only had to increase my assumed inert fraction a little
bit to achieve an airframe-only inert fraction that I considered to be
believable. Even so, the payload fraction is about twice the
payload fraction of the two-stage expendable chemical vehicle, and almost 4 times the payload fraction of
the one-stage expendable chemical vehicle.
And the nuclear one-stage vehicle is entirely reusable, but if <u>and only if</u> you can accept
returning its engine to Earth!<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhxnpYrmJSrlDWYjtTxV2gyuCFwFhWuXCi09H-xRp7bMm0LxLmFy4y3B0vDqBas4KUqgOaOE0QWRHAiSe0nGjVDew-mSuhV6EylPR4RW1XS0H-nJw15pnQS0VczAxgA8IU6_eQFMPl0YxW7ZmtS2DMYl0XyfOXVzhk6aUoPGpyIfrYIlTjWuKTR8FYYLVdq/s1466/F6%201%20stage%20nuke.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="533" data-original-width="1466" height="145" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhxnpYrmJSrlDWYjtTxV2gyuCFwFhWuXCi09H-xRp7bMm0LxLmFy4y3B0vDqBas4KUqgOaOE0QWRHAiSe0nGjVDew-mSuhV6EylPR4RW1XS0H-nJw15pnQS0VczAxgA8IU6_eQFMPl0YxW7ZmtS2DMYl0XyfOXVzhk6aUoPGpyIfrYIlTjWuKTR8FYYLVdq/w400-h145/F6%201%20stage%20nuke.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 6 – Spreadsheet Image for the Single-Stage Nuclear
Vehicle<o:p></o:p></p><p class="MsoNormal"> <b><i>Final
Remarks<o:p></o:p></i></b></p><p class="MsoNormal">For the nearer term,
using only well-developed,
ready-to-apply technologies, the
highest payload fraction option is the two-stage vehicle, which can readily adapt the designs of its
two stages to the different circumstances of ascent out of the atmosphere, and acceleration exo-atmospheric and nearly
horizontal to orbital speed. Making its
first stage reusable would not cost that much payload fraction. <o:p></o:p></p><p class="MsoNormal">Trying to do this,
even if expendable, as a
one-stage vehicle with chemical propulsion, is unlikely to provide a payload fraction high
enough to actually pay off. It will
likely underperform the two-stage expendable in terms of payload fraction, no matter what propellants might be used. And it will be heavier at liftoff under any
conceivable circumstances, for the same
payload. Thus it will cost more.<o:p></o:p></p><p class="MsoNormal">Longer term, a fully
reusable one-stage vehicle of even higher payload fraction than the two-stage
expendable chemical vehicle, might be
feasible with some form of nuclear thermal propulsion that performs only
slightly better than NERVA. Key to its
viability will be the acceptability of returning and landing with that engine
aboard.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEihMJntmHYvD8gyupx4SCE1ZzRbQOuhntgkhGxBgR9AZe4_gYbOMIDRbpbDZOZCkRfp_RyYbg61P0qmMMX9h3MsT5vJbyAd_6A_GUYz0esfyx4_cTgIp8RDhHOf8gGmsLczFJkQ_55Pihyphenhyphen4UXJa9_y-bKLV-voWcRvlh-X50WNyGNuHgdA-0evgUoquwCiD/s973/F7%20nuke%20design.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="608" data-original-width="973" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEihMJntmHYvD8gyupx4SCE1ZzRbQOuhntgkhGxBgR9AZe4_gYbOMIDRbpbDZOZCkRfp_RyYbg61P0qmMMX9h3MsT5vJbyAd_6A_GUYz0esfyx4_cTgIp8RDhHOf8gGmsLczFJkQ_55Pihyphenhyphen4UXJa9_y-bKLV-voWcRvlh-X50WNyGNuHgdA-0evgUoquwCiD/w400-h250/F7%20nuke%20design.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 7 – The Nuclear One-Stage Vehicle Concept<o:p></o:p></p><p class="MsoNormal"><b><u>Update 3-6-2024</u>:
<o:p></o:p></b></p><p class="MsoNormal">I went ahead and looked more closely at the engines for the
two-stage vehicle. These would be LOX-LH2
in the second stage with vacuum bell designs,
and LOX-RP1 in the first stage,
with something suitable as a sea level bell design. <b><i>Neither would push the state of the art
the way the LOX-LCH4 engines must do, in
the one-stage vehicle.</i></b> I used
very modest modern-technology characteristics for the engines of both
stages: 2500 psia max Pc, with only a P-TDR = 2.5, and a dumped bleed fraction BF = 0.02. They use otherwise the same 18-8<sup>o</sup>
bell profile and C<sub>D</sub> = 0.995.<o:p></o:p></p><p class="MsoNormal">As <b>Figure 8</b> shows,
the traditional sea level design with perfect expansion to sea level
pressure from max Pc, shows an
ascent-averaged Isp shortfall relative to what I wanted for the first
stage. But when I used the “compromise design”
approach (see <b>Figure 9</b>) to size those engines, trading away unseparated sea level operation
at min-throttle setting, for more
expansion ratio and higher vacuum and ascent-averaged Isp values, that ascent-averaged Isp exceeded the
assumptions used for roughing out the first stage. <o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal">Elsewhere, I had
looked at vacuum designs for LOX-LH2 engines,
sized to arbitrary expansion ratios of A/A* = 100, 150,
and 200, with those same modest
modern-technology characteristics. The
min expansion version (<b>Figure 10</b>) gets you the smallest physical length
and exit diameter, and its vacuum Isp
substantially exceeded what I assumed for the rough-sizing of the second
stage. So, I revisited the vehicle rough-out with a
somewhat-higher second stage Isp (<b>Figure 11, </b>blue edits). That increased the payload fraction, and reduced the launch weight, both acting to lower costs. <o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg8wU3Smana6d_c1HKijoWeAY_ZZ91pReaCT6nWLDGrXFLIsf_N0Ut173A67kBV_pVnO98A8BqMfkQaJkjGosA2MkBwvLIlNnCXAOBMIvJ7L-M80Jbcs6aPxPcoJzKonXz24nzNpirwNmmccICJlYOdvOSGr8cIFkc2TxOX-kKlDlZ6T3wxG4p3Ar-glDQd/s977/sea%20level%20kerolox.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="611" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg8wU3Smana6d_c1HKijoWeAY_ZZ91pReaCT6nWLDGrXFLIsf_N0Ut173A67kBV_pVnO98A8BqMfkQaJkjGosA2MkBwvLIlNnCXAOBMIvJ7L-M80Jbcs6aPxPcoJzKonXz24nzNpirwNmmccICJlYOdvOSGr8cIFkc2TxOX-kKlDlZ6T3wxG4p3Ar-glDQd/w400-h250/sea%20level%20kerolox.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 8 – Traditional Sea Level Sizing Falls Short of Desired
Ascent-Averaged Isp = 330 s<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiBQwRA7K6QnriGmTY70LN5GfPm2mBKAjtW3FRWNWepvXd-C5yupjifYlGddNCvpxrOLXRpu1bbSwVzK1bBTUkz-NdYH9VHXAJpe7QvSLmXJmDT8izPHV0ya5k8fR8hXoKBkjf0zPutdMSOBfPwoHVFENhkSyK8mV5biWhiri4wcpIY4YMPGMHRSR6ZMopT/s977/underexpanded%20kerolox.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="613" data-original-width="977" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiBQwRA7K6QnriGmTY70LN5GfPm2mBKAjtW3FRWNWepvXd-C5yupjifYlGddNCvpxrOLXRpu1bbSwVzK1bBTUkz-NdYH9VHXAJpe7QvSLmXJmDT8izPHV0ya5k8fR8hXoKBkjf0zPutdMSOBfPwoHVFENhkSyK8mV5biWhiri4wcpIY4YMPGMHRSR6ZMopT/w400-h251/underexpanded%20kerolox.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 9 – Sea Level “Compromise” Design Exceeds Desired Ascent-Averaged
Isp = 330 s<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj-Nz-DmqnASMLUSQnXYojzhfq1JfwALnxd-EAsBUgf5T_dfKnTw7PGvFMtF-DI5dEvO50R7tMrrH7yidEPtdCzGYJrPrHXVHRrTx0xmSD_ZImTFDGiseKY9sKm3JpJOUvbwwqsOL2X07oStix-Wu01YJJFbbX6wvAUKjTPX99AtxXXtmQPwq14XDz-OhfF/s973/vacuum%20100.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="973" height="251" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj-Nz-DmqnASMLUSQnXYojzhfq1JfwALnxd-EAsBUgf5T_dfKnTw7PGvFMtF-DI5dEvO50R7tMrrH7yidEPtdCzGYJrPrHXVHRrTx0xmSD_ZImTFDGiseKY9sKm3JpJOUvbwwqsOL2X07oStix-Wu01YJJFbbX6wvAUKjTPX99AtxXXtmQPwq14XDz-OhfF/w400-h251/vacuum%20100.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 10 – Vacuum Design at A/A* = 100 Substantially
Exceeds Desired Vacuum Isp = 450 s<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh9k2MYYKGfqg8SQREEpk57qF4CbM9xeYa7AY2xgUuBUm3mofzVzFEIbFVmT0Dg1R3Yo0BCly07JyaN60PgYARfklSadx_fgwa8qrYfE_vT3j9ulNdlPE9j76sB7mFjJmU2PIW6U08jITEkt-F6lcUv1jV-vPOMan1mA_-uhXVnPEKK3rpS0wevEWVHqF1Q/s977/veh%20rough%203.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh9k2MYYKGfqg8SQREEpk57qF4CbM9xeYa7AY2xgUuBUm3mofzVzFEIbFVmT0Dg1R3Yo0BCly07JyaN60PgYARfklSadx_fgwa8qrYfE_vT3j9ulNdlPE9j76sB7mFjJmU2PIW6U08jITEkt-F6lcUv1jV-vPOMan1mA_-uhXVnPEKK3rpS0wevEWVHqF1Q/w400-h250/veh%20rough%203.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 11 – Revised Vehicle Rough-Outs Show 2-Stage To Be
Even Slightly Better<o:p></o:p></p><p class="MsoNormal"><b><u>Update 3-7-2024</u>:
<o:p></o:p></b></p><p class="MsoNormal">
</p><p class="MsoNormal">The question came up of whether I demanded enough dV of the
vehicles? I had added 10% to the 7.7
km/s orbital velocity for 8.5 km/s. Here
is what the size-out produces with 20% added, for 9.2 km/s.
The “best” engines and nozzles that I found earlier were retained just as
they were revised. <u>Only the velocity
requirement was increased</u>. I put all
the increased burden on the first stage of the two-stage vehicle, precisely because it has the lower Isp, as a worst case. See <b>Figure 12</b>. <o:p></o:p></p><p> <a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhou8tE-Mn9VnKrxm0ee6hbjOYJFOR536lds9aWlqgkPGE-SxZycKF78OTVgj9gZz3HQwOWPEOcCeN-isGATLWD7EidkepEEPyCURe9M1VAdEdSXkUWbhYW_elpfBh5qI9akH6dClP9Ulp7Q_OWjYAxUXvkK6IN1DpOxzryk5Pk43nJQe-iqMXMfp2GvQ2e/s977/revised%20launch%20reqmnts.png" style="margin-left: 1em; margin-right: 1em; text-align: center;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhou8tE-Mn9VnKrxm0ee6hbjOYJFOR536lds9aWlqgkPGE-SxZycKF78OTVgj9gZz3HQwOWPEOcCeN-isGATLWD7EidkepEEPyCURe9M1VAdEdSXkUWbhYW_elpfBh5qI9akH6dClP9Ulp7Q_OWjYAxUXvkK6IN1DpOxzryk5Pk43nJQe-iqMXMfp2GvQ2e/w400-h250/revised%20launch%20reqmnts.png" width="400" /></a></p><div class="separator" style="clear: both; text-align: center;"><br /></div><p></p><p class="MsoNormal">Figure 12 – Revised Launch Requirements for Higher dV Values
for Rough-Out<o:p></o:p></p><p class="MsoNormal">The resulting vehicle size-outs show larger vehicles and
lower payload fractions, to be sure <u>exactly
as expected</u>! However, the SSTO is now worse by about a factor of
3, not just 2, than the two stage vehicle. Both were considered to be expendables for
this, as before. See <b>Figure 13</b>. The engine count is getting to be something to
worry about, as well. That cluster has to fit behind a slender
tankage set. If you make the tanks fatter
and shorter to “cover” the cluster, you
are no longer “long and slender”, and
that increases your drag loss. <o:p></o:p></p><p class="MsoNormal">One thing readers should consider is the requirement for
adequate kinematics right off the launch pad.
<b><i>You need half a gee or more, of effective net acceleration beyond
gravity, to be efficient, and not spend most of your propellant just
climbing the first few thousand feet.</i></b>
For an Earth launch, that’s an <b><i>ignition
thrust/weight of 1.5 or higher.</i></b> <o:p></o:p></p><p class="MsoNormal">In the real world, you
can use more and/or bigger engines to achieve this, or you can add some solids (always of much
higher frontal thrust density than a cluster of liquids). I chose to just use more and bigger
engines. Why complicate the study?<o:p></o:p></p><p class="MsoNormal">This is a <u>very strong effect</u>, almost to the point of being overwhelming! It is precisely why vehicles with low launch
thrust/weight also have historically had low payload fractions. The poor acceleration kinematics drastically
raise the gravity loss, making the dV
requirement effectively much larger, and
THAT lowers payload fraction rapidly. To
be “efficient”, you really have to scoot
off the pad! And THAT is exactly what I
enforced in this study!<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjI3AD6a-bzcgWF-5U3DIrYRgTn45yDgRhpQDzw7rGD4ODFSVSO42dTEtgv3SwOD_NO0IPM6ym5mqApjCAWFxn8GyEFSXjD_OFAX_aHhEW_g7_BSoauFlBuMkzg-ZFNdg-lyVzd6yXNbLnPwVvHzynWDXwdOQCtg6z-N040HRKS5OanlGfg-i5er1UVI2uC/s977/F2%20veh%20rough.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="977" height="250" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjI3AD6a-bzcgWF-5U3DIrYRgTn45yDgRhpQDzw7rGD4ODFSVSO42dTEtgv3SwOD_NO0IPM6ym5mqApjCAWFxn8GyEFSXjD_OFAX_aHhEW_g7_BSoauFlBuMkzg-ZFNdg-lyVzd6yXNbLnPwVvHzynWDXwdOQCtg6z-N040HRKS5OanlGfg-i5er1UVI2uC/w400-h250/F2%20veh%20rough.png" width="400" /></a></div><p class="MsoNormal">Figure 13 – Revised Rough-Out Results for Higher dV Requirement<o:p></o:p></p><p class="MsoNormal">---------- </p><p class="MsoNormal">I should probably have followed my own recommendations and used the surface circular orbit speed of 7.9 km/s, and not the speed at orbit altitude 7.7 km/s, as the "ideal dV" to be factored up for gravity and drag. But it really doesn't matter very much when doing a comparison analysis. The factor would be 1.1 if one assumes 5% each for gravity and drag losses. It is 1.20 if instead one assumes 10% each for gravity and drag. 7.7*1.1 = 8.5, some 0.8 loss to cover, while 7.7*1.2 = 9.2, some 1.5 loss to cover. If instead you use 7.9 km/s, the numbers are only slightly different: 7.9*1.1 = 8.7, for 0.8 loss, and 7.9*1.2 = 9.5, for some 1.6 loss. </p><p class="MsoNormal">More important is arriving on orbit with something left to support doing rendezvous, plus some sort of controlled de-orbit burn. The former is likely on the order of 0.3-0.5 km/s, and the latter is about 0.1 km/s, for about an extra 0.5 km/s. You add those unfactored to the total effective launch dV. That would be around 8.5+0.5 = 9.0, or 9.2+0.5 = 9.7. I did NOT include anything like that in the dV requirements, because I was only looking for relative trends. </p><p class="MsoNormal">And those relative trends say the single-stage-to-orbit (SSTO) does factor 2-to-3 worse in terms of payload fraction and launch weight than the two-stage design. That's for both designs being clean and slender for low drag loss, and neither pushing the state-of-the-art on structure technologies (the fixed 5% inert in a loaded stage). The 2-stage does not push the state-of-the-art on its engine technologies, but the SSTO has to. SpaceX has already had its troubles with that, in its own LOX-LCH4 engines.</p><p class="MsoNormal">If you go to LOX-LH2 to improve past the Isp of LOX-LCH4 for the SSTO, you will end up having to push the state-of-the-art on your structure technologies as well as your engine technologies. And it may no longer qualify as "clean-and-slender, so the drag factor may increase, too. </p><p class="MsoNormal">If you want something easily and less-expensive to develop, then don't push the state-of-the-art. If you do, you will have higher development costs to amortize. Everything is acting in the wrong direction on costs, with a chemical SSTO. </p><br /><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com6tag:blogger.com,1999:blog-2675974463524895416.post-6174607067651422182024-02-25T16:41:00.002-06:002024-02-26T11:46:03.322-06:00Tricky Landing<p>The “Odysseus” robot lander created by Intuitive Machines
seems to have landed successfully on the moon,
although reports say it is on its side rather than upright. Details as of yet are quite sparse, but depending upon whose reports you
read, it would appear the lander had a
non-trivial and unintended horizontal speed at touchdown. Odds are,
it was also tilted a bit in the direction of that horizontal
motion. It seems to have “tripped” on
one of its legs being somehow obstructed,
overturning the lander as it touched down.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">This is always going to be a serious problem for a robot
moon lander, as opposed to a manned
craft. It is still difficult-to-impossible
to program a robot to do what a human pilot can do, and robot vision is still nowhere near as
good as human vision. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">Consider what Neil Armstrong had to do, landing the Apollo 11 lunar module. The computer-controlled trajectory was taking
the vehicle into a tightly-packed field of multiple boulders as big as houses: a guaranteed fatal crash! Armstrong had to take manual control, stop the descent into an unplanned
hover, and then direct that hover toward
a clear landing site nearby. The
depletion of his rocket fuel was but a single handful of seconds away at engine
shutdown. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">The rule-of-thumb stability criterion used for all
successful landing leg-equipped craft on the moon and Mars, is that the span between foot pads at least
equals the height of the craft center of gravity, and preferably exceeds it. The Odysseus
lander only just barely met this, and
the also-recent Japanese lander did not meet it, and was photographed upside-down after its landing!</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">It is not yet known what “tripped” the Odysseus lander, but the odds favor either a leg striking a
fixed rock or similar obstruction, or
else a landing pad digging into the surface dirt. The “fix” for this is two-fold: (1) increase the pad span to
center-of-gravity height ratio significantly,
and (2) hinge the pads on the ends of the legs, and spring-load them to tilt upward toward
the radial-outward direction at each leg.
The first decreases the net overturning torque of a “tripping”
incident, and the second acts to prevent
a pad from digging-in. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">Seems “intuitive” to me.
Maybe we old farts still have things to contribute, after all. </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhxOhXhxYqUmRKRdek9A93enap0-2bwqplvbDgyn595vIsVo9WjerwbS8ue3UAfG-P00q5OGrtycreVDtrflAjUKEEiw8GfmOG38CL-5oGbrkgLii9s4heFuHvHSMVN9Um2R7N7HOehOLDzJGSUq67ThzP7_M2gbCX16c4PbVS6X2Onha7pDR1EgPG8GEUK/s1007/tricky%20landing.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="616" data-original-width="1007" height="245" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhxOhXhxYqUmRKRdek9A93enap0-2bwqplvbDgyn595vIsVo9WjerwbS8ue3UAfG-P00q5OGrtycreVDtrflAjUKEEiw8GfmOG38CL-5oGbrkgLii9s4heFuHvHSMVN9Um2R7N7HOehOLDzJGSUq67ThzP7_M2gbCX16c4PbVS6X2Onha7pDR1EgPG8GEUK/w400-h245/tricky%20landing.png" width="400" /></a></div><br /><p class="MsoNormal"><b><u>Update 2-26-2024</u>:</b> The quote below, from a PBS NewsHour story published this date,
confirms what I hypothesized about the
Odysseus lander. It apparently will
cease operations tomorrow for lack of sunlight on its solar panels, according to the story. Highlighting is mine. <o:p></o:p></p><p class="MsoNormal"><span style="background: white;">“The lander,
named Odysseus, is the </span><a href="https://www.pbs.org/newshour/science/private-lunar-lander-circling-moon-for-first-u-s-touchdown-in-half-century" style="box-shadow: rgb(214, 223, 232) 0px -1px 0px inset; font-feature-settings: inherit; font-kerning: inherit; font-optical-sizing: inherit; font-stretch: inherit; font-variant-alternates: inherit; font-variant-east-asian: inherit; font-variant-numeric: inherit; font-variant-position: inherit; font-variation-settings: inherit; line-height: inherit; transition: color 0.2s ease 0s, box-shadow 0.2s ease 0s, -webkit-box-shadow 0.2s ease 0s;"><b><span style="background: white; border: 1pt none windowtext; color: black; padding: 0in; text-decoration-line: none;">first U.S. spacecraft to land on the moon</span></b></a><span style="background: white;"> in more than 50 years, carrying experiments for NASA,
the main sponsor. <span style="background: yellow; mso-highlight: yellow;">But it
came in too fast last Thursday and the foot of one of its six legs caught on
the surface, causing it to tumble over,</span> according to company officials.”</span><span style="background: white;"><o:p></o:p></span></p><p class="MsoNormal"><span style="background: white;">What I found
about the Japanese “SLIM” lander is enlightening, although it is still unclear just exactly how
it ended up on its nose. It was supposed
to hover and then tip over onto its side,
with its landing legs extending out that side. In the long dimension, pad span exceeds cg height, meeting the criterion. From side-to-side, it does not meet the criterion, a very real risk. </span><span style="background: white; mso-bidi-font-family: Calibri; mso-bidi-theme-font: minor-latin;"><o:p></o:p></span></p><p class="MsoNormal"><span style="background: white;">But it did
not fall over to one side, somehow it instead
went tumbling end-over-end, which is the
only way it could have ended up on its nose!
Some stories mention a problem with a main thruster (there were two in
its bottom). Those should be “off” during
the actual tip-over-and-landing because they are too powerful, so any main thruster-related hypothesis would
have to have occurred before that process.</span><span style="background: white; mso-bidi-font-family: Calibri; mso-bidi-theme-font: minor-latin;"><o:p></o:p></span></p><p class="MsoNormal">
</p><p class="MsoNormal"><span style="background: white;">One possible
main thruster-related hypothesis is that it may have experienced suddenly-asymmetric
thrust just as it approached hover for final tip-over-and-landing. If so,
that could have sent it tumbling end-over-end while still aloft, instead of hovering into a controlled
tip-over. There should be marks in the regolith if that hypothesis is
true, marks where it hit while already tumbling
end-over-end. In the low gravity, it would have continued to bounce
end-over-end after hitting the surface.
It just happened to quit bouncing,
while on its nose.
Improbable, but possible. Still, only a hypothesis. </span><o:p></o:p></p><p class="MsoNormal"><o:p></o:p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com2tag:blogger.com,1999:blog-2675974463524895416.post-50111003853273357732024-02-12T08:59:00.000-06:002024-02-12T08:59:34.769-06:00GW’s Ramjet Book Is Now Available!<p>This has been a long time coming, because I originally finished writing the
book back in 2017. I offered it to a
technical publisher, who took over 2
years to decline publishing it as a hardcopy,
hardcover book. It then
languished as a back-burner item, while
I figured out how to really do this myself,
in between more pressing obligations.
But I knew that I really did need to get this book “out there”, because I am an old retired person, and so I will not be around for that many
years yet to come. I’d rather this
knowledge and experience to not die with me.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal"><b><i>This is not your usual academic tome.</i></b><span style="mso-spacerun: yes;"> </span>It is more of a very extensive “how-to”
compendium of the <u>things that actually worked for me and my colleagues</u>,<span style="mso-spacerun: yes;"> </span><u>while actually doing real ramjet work</u>.<span style="mso-spacerun: yes;"> </span>The scope is subsonic combustion ramjet,<span style="mso-spacerun: yes;"> </span>not combined cycles,<span style="mso-spacerun: yes;"> </span>and not supersonic combustion ramjet
(scramjet).<span style="mso-spacerun: yes;"> </span>It covers mainly
liquid-fueled and solid gas generator-fed ramjets.<span style="mso-spacerun: yes;"> </span>It bears about the same relationship to
ramjet engineering, <span style="mso-spacerun: yes;"> </span>that Sighard F.
Hoerner’s self-published books, <span style="mso-spacerun: yes;"> </span>“Fluid
Dynamic Drag” and “Fluid Dynamic Lift”, <span style="mso-spacerun: yes;"> </span>bore to aerodynamical engineering.<span style="mso-spacerun: yes;"> </span>No real publisher wanted to publish those
books,<span style="mso-spacerun: yes;"> </span>either,<span style="mso-spacerun: yes;"> </span>but scads of people found them very useful
anyway!<span style="mso-spacerun: yes;"> </span>I hope you find my book
useful.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal"><b><i>I now have an initial solution:<span style="mso-spacerun: yes;"> </span>I can literally email the book as pdf files
to those who want to buy it.</i></b><span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">It exists as some 27 pdf files:<span style="mso-spacerun: yes;"> </span>one for the “up-front” stuff,<span style="mso-spacerun: yes;"> </span>one each for all 22 chapters,<span style="mso-spacerun: yes;"> </span>and one each for all 4 appendices.<span style="mso-spacerun: yes;"> </span>Each chapter has its own page numbering,<span style="mso-spacerun: yes;"> </span>its own figure numbering,<span style="mso-spacerun: yes;"> </span>and its own reference list.<span style="mso-spacerun: yes;"> </span>The “up-front” stuff includes a
foreword,<span style="mso-spacerun: yes;"> </span>biographical data,<span style="mso-spacerun: yes;"> </span>a table of contents,<span style="mso-spacerun: yes;"> </span>and (quite uniquely !!) <u>another table of
contents <b>with a paragraph indicating content</b> for each chapter</u>!<o:p></o:p></p>
<p class="MsoNormal"><b><span style="background: yellow; mso-highlight: yellow;">If
you want to buy the book,<span style="mso-spacerun: yes;"> </span>just contact
me,<span style="mso-spacerun: yes;"> </span>email is best!<span style="mso-spacerun: yes;"> </span>My email is </span></b><a href="mailto:gwj5886@gmail.com"><b><span style="background: yellow; mso-highlight: yellow;">gwj5886@gmail.com</span></b></a><b><span style="background: yellow; mso-highlight: yellow;">.<span style="mso-spacerun: yes;"> </span>(I will need
your email to send the files,<span style="mso-spacerun: yes;"> </span>in any
event.)<span style="mso-spacerun: yes;"> </span>I will give you my physical
address,<span style="mso-spacerun: yes;"> </span>to which you can send the
purchase amount by check or money order.<span style="mso-spacerun: yes;">
</span><o:p></o:p></span></b></p>
<p class="MsoNormal"><b><span style="background: yellow; mso-highlight: yellow;">When
I receive it,<span style="mso-spacerun: yes;"> </span>I will email the files to
you in multiple emails,<span style="mso-spacerun: yes;"> </span>since there are
so many and they are large.<span style="mso-spacerun: yes;"> </span>Plus,<span style="mso-spacerun: yes;"> </span><u>I will follow up</u> to make sure you get
them all!<span style="mso-spacerun: yes;"> </span><o:p></o:p></span></b></p>
<p class="MsoNormal"><b><span style="background: yellow; mso-highlight: yellow;">Base
price is $100 per copy.<span style="mso-spacerun: yes;"> </span>Out here
rural,<span style="mso-spacerun: yes;"> </span>the Texas sales tax is only 6.25%
(and I have a Texas sales tax certificate),<span style="mso-spacerun: yes;">
</span>so the sales tax amount is $6.25.<span style="mso-spacerun: yes;">
</span>That puts the total purchase amount at $106.25,<span style="mso-spacerun: yes;"> </span>turn-key.</span><span style="mso-spacerun: yes;"> </span><o:p></o:p></b></p>
<p class="MsoNormal">That’s how I need to do this for now.<span style="mso-spacerun: yes;"> </span>Soon I hope to be able to take credit cards by
voice over the phone,<span style="mso-spacerun: yes;"> </span>which would speed
the process up for you.<span style="mso-spacerun: yes;"> </span>But that is not
ready yet.<span style="mso-spacerun: yes;"> </span>Watch this space for
updates,<span style="mso-spacerun: yes;"> </span>I will add that capability
soon.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Eventually,<span style="mso-spacerun: yes;"> </span>if I can
find qualified help,<span style="mso-spacerun: yes;"> </span>I hope to set up another
site that automates the payment and send-out processes.<span style="mso-spacerun: yes;"> </span>But that is for the future.<o:p></o:p></p>
<p class="MsoNormal">-----------------<o:p></o:p></p>
<p class="MsoNormal">As examples of my real-world experience,<span style="mso-spacerun: yes;"> </span>I have included some pictures here.<span style="mso-spacerun: yes;"> </span>I hope these help inspire in you enough
confidence that I really know what I am talking about,<span style="mso-spacerun: yes;"> </span>so that you will be more inclined to buy the
book. <o:p></o:p></p>
<p class="MsoNormal">Figure 1 below shows the hybridized ground test hardware
that I used to great advantage, <span style="mso-spacerun: yes;"> </span>doing
ramjet tests on the ground in a direct-connect facility, <span style="mso-spacerun: yes;"> </span>long ago.<span style="mso-spacerun: yes;">
</span>It coupled a heavyweight lab motor as a short-burn solid gas
generator,<span style="mso-spacerun: yes;"> </span>to a flight-weight combustor
and inlets.<span style="mso-spacerun: yes;"> </span>This hardware was extremely
effective for testing experimental fuel propellants,<span style="mso-spacerun: yes;"> </span>experimental combustor insulations,<span style="mso-spacerun: yes;"> </span>and one experimental fuel flow rate control
approach. <o:p></o:p></p>
<p class="MsoNormal">Figure 2 below illustrates what the exhaust plume looked
like for one of these tests,<span style="mso-spacerun: yes;"> </span>conducted
“open-air nozzle”,<span style="mso-spacerun: yes;"> </span>with <u>both</u> an
experimental fuel <u>and</u> an experimental combustor insulation,<span style="mso-spacerun: yes;"> </span>back in 1991.<span style="mso-spacerun: yes;">
</span>Most of the sparkler streaks were from the insulation,<span style="mso-spacerun: yes;"> </span>not the fuel!<span style="mso-spacerun: yes;">
</span><o:p></o:p></p>
<p class="MsoNormal">Believe it or not,<span style="mso-spacerun: yes;">
</span>this particular test was the first time anybody ever burned
high-percentage boron <u>efficiently</u> in a ramjet!<span style="mso-spacerun: yes;"> </span>See the smoke-free clarity of the plume
downstream of the fire as proof.<span style="mso-spacerun: yes;"> </span>Even
the metal oxide smoke is barely visible downstream.<span style="mso-spacerun: yes;"> </span>However,<span style="mso-spacerun: yes;">
</span>the incandescent glare from it (and some soot) is part of what makes the
tailpipe flame opaquely brilliant.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Figure 3 below shows a modern cutaway display model of the
Russian surface-to-air missile known to NATO as the SA-6 “Gainful”.<span style="mso-spacerun: yes;"> </span>It was a solid gas generator fed ramjet,<span style="mso-spacerun: yes;"> </span>with an “integral booster”,<span style="mso-spacerun: yes;"> </span>meaning the booster rocket was housed inside
the combustor itself,<span style="mso-spacerun: yes;"> </span>not a staged-off
item.<span style="mso-spacerun: yes;"> </span>As a young engineer, <span style="mso-spacerun: yes;"> </span>I was the lone engineer among 3 propellant
chemists who did the actual exploitation work on this foreign technology. <span style="mso-spacerun: yes;"> </span>We duplicated the solid fuel-rich propellant
and its processing,<span style="mso-spacerun: yes;"> </span>tested it static and
with air,<span style="mso-spacerun: yes;"> </span>and I put together a computer
trajectory model of the system,<span style="mso-spacerun: yes;"> </span>which
matched performance seen on the battlefield.<span style="mso-spacerun: yes;">
</span>This knowledge used to be classified,<span style="mso-spacerun: yes;">
</span>but no longer (not with public display models). <o:p></o:p></p>
<p class="MsoNormal">Figure 4 below is a two-view picture of an ASALM-PTV being
launched from an A-7 Corsair-2.<span style="mso-spacerun: yes;"> </span>It was a
liquid-fueled ramjet with an integral booster.<span style="mso-spacerun: yes;">
</span>I worked on ASALM,<span style="mso-spacerun: yes;"> </span>which was a
prototype for a high-altitude supersonic ramjet-powered cruise missile that
(unfortunately in my opinion) never proceeded to operational status,<span style="mso-spacerun: yes;"> </span>because of treaty limitations.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">It was flight tested 7 times,<span style="mso-spacerun: yes;"> </span>and met or exceeded all objectives on 6 of
those.<span style="mso-spacerun: yes;"> </span>On the first of those tests,<span style="mso-spacerun: yes;"> </span>it accidentally went hypersonic due to an
assembly error in its fuel throttle controls.<span style="mso-spacerun: yes;">
</span>Way back in 1980,<span style="mso-spacerun: yes;"> </span>this thing
accelerated <u>in ramjet</u>,<span style="mso-spacerun: yes;"> </span><u>at low
altitude</u>,<span style="mso-spacerun: yes;"> </span>to about Mach 6!<span style="mso-spacerun: yes;"> </span>A different picture of it (not here) on the same
airplane,<span style="mso-spacerun: yes;"> </span>I consider to be a sort of
“family portrait”.<span style="mso-spacerun: yes;"> </span>I worked on
ASALM,<span style="mso-spacerun: yes;"> </span>and my father was lead
engineering designer for the A-7.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgsUFbKtAsO5RFWekhHCg56FNweAtyTs4jh4rXZHPisjOMsFTuppG_mUREIhpw9E2N0d0h4uObNYUPEdJMR17XgAJEY180B7xv3oxfmUd0Mh7PuKiBYtyW4TBKtotVCbzHwFjk5jANCZMZW2GgZqZ5uEMKD-DensF6r41r3D7vP07eEpPibJr2A9RKm4sBX/s2968/test%20hardware.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="2372" data-original-width="2968" height="320" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgsUFbKtAsO5RFWekhHCg56FNweAtyTs4jh4rXZHPisjOMsFTuppG_mUREIhpw9E2N0d0h4uObNYUPEdJMR17XgAJEY180B7xv3oxfmUd0Mh7PuKiBYtyW4TBKtotVCbzHwFjk5jANCZMZW2GgZqZ5uEMKD-DensF6r41r3D7vP07eEpPibJr2A9RKm4sBX/w400-h320/test%20hardware.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure1 – Hybridized Ground Test Hardware<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_g-eGU91yA711QVFfkvK2iEFAqJGQQFOrbffLiCn1LsT4lbv2M224RBkf_dLtSOBaB1Jgyxm4BV9_U-8ULR0AJfiLjwCMIjjKjnkuBVq22CVAnlr0JKxAXe3Bw2cg6SX7ltynP_pyNiKslvsGqvYFQ4jAhH51m6OlVUBlIIaeJUl2W2bNjmclQsS6mReP/s2928/Bti%20test.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="2352" data-original-width="2928" height="321" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_g-eGU91yA711QVFfkvK2iEFAqJGQQFOrbffLiCn1LsT4lbv2M224RBkf_dLtSOBaB1Jgyxm4BV9_U-8ULR0AJfiLjwCMIjjKjnkuBVq22CVAnlr0JKxAXe3Bw2cg6SX7ltynP_pyNiKslvsGqvYFQ4jAhH51m6OlVUBlIIaeJUl2W2bNjmclQsS6mReP/w400-h321/Bti%20test.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2 – Typical Experimental Open-Air Nozzle Ramjet Test<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEitYnEo0pa3hRcIQr-gblFS_hSN8y1azB-sCx8XLV_s0c5TtBVCMUsIvvvWMIijvBagpJZJ4uIfgfwR5T-AQjqwJJzXvO2UZ8wIYkAHrPCTsUHxD3ba1In3LJD4cXjrzJncR30hJ144vMDD3diRjKZxeXoxIYXzDGcIwXf4kLo-JtYsx902TIHAan4O3JcV/s975/exploited%20engine%20for%20SA-6.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="439" data-original-width="975" height="180" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEitYnEo0pa3hRcIQr-gblFS_hSN8y1azB-sCx8XLV_s0c5TtBVCMUsIvvvWMIijvBagpJZJ4uIfgfwR5T-AQjqwJJzXvO2UZ8wIYkAHrPCTsUHxD3ba1In3LJD4cXjrzJncR30hJ144vMDD3diRjKZxeXoxIYXzDGcIwXf4kLo-JtYsx902TIHAan4O3JcV/w400-h180/exploited%20engine%20for%20SA-6.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 3 – Modern Cutaway Display Article of the SA-6<o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhAMApt3EpAb98p5XOh8rp0heWFfEpqPvifIx-Dt7lZRdCUEbOQkBI0pMmy1sJWpUZqA1JFLouowAyjKS1WPQBqodgzqaptNRU9wVTnadG4emar_cRLumsEw5Fi-xdJzJ7_hKbguB4qcC3zpC5wWK9jYWLhWaZ5YDjbLusBp6_SYFshmsMLl3R_d4mZwdFX/s368/ASALM-PTV.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="137" data-original-width="368" height="149" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhAMApt3EpAb98p5XOh8rp0heWFfEpqPvifIx-Dt7lZRdCUEbOQkBI0pMmy1sJWpUZqA1JFLouowAyjKS1WPQBqodgzqaptNRU9wVTnadG4emar_cRLumsEw5Fi-xdJzJ7_hKbguB4qcC3zpC5wWK9jYWLhWaZ5YDjbLusBp6_SYFshmsMLl3R_d4mZwdFX/w400-h149/ASALM-PTV.jpg" width="400" /></a></div><p></p><p class="MsoNormal">Figure 4 – Two-View of ASALM-PTV Launched From A-7 Corsair-2<o:p></o:p></p><p class="MsoNormal"><span style="mso-spacerun: yes;"><br /></span></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com1tag:blogger.com,1999:blog-2675974463524895416.post-44685267997628157502024-02-01T15:54:00.000-06:002024-02-01T15:54:31.654-06:00Swatting at Proxies is Pointless<p>There has been a lot of violence and danger in the Middle
East in recent decades. Most, but not quite all of it, is summarized in Figure 1. This article is about things since about
1980. For earlier history, see “Israel vs Hamas: It Is Worse Than You Think”, posted 29 December, 2023,
on this same site. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">I drew that figure on 22 January.<span style="mso-spacerun: yes;"> </span>Since I drew the figure,<span style="mso-spacerun: yes;"> </span>there has been an attack by Iranian proxies
on a US base in Jordan,<span style="mso-spacerun: yes;"> </span>killing 3 of our
troops.<span style="mso-spacerun: yes;"> </span>However,<span style="mso-spacerun: yes;"> </span>that still fits the same pattern shown!<o:p></o:p></p>
<p class="MsoNormal">Note that Western nations in general,<span style="mso-spacerun: yes;"> </span>and Israel in particular,<span style="mso-spacerun: yes;"> </span>have been combatting the proxies listed since
about 1980.<span style="mso-spacerun: yes;"> </span>There is however a common
thread to all of that listed evil,<span style="mso-spacerun: yes;"> </span>and
that common thread is shown in the figure to be Iran. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgbH6i1tjAWCE0A1K3uud_J4NXP51n1xkPdikthYF6x171wUc7Fy5HHJZBRVZRfT-QwCdwR2IBR_U-BIRmWgB-hkehm5lpinAEC5XjVzm8Ktk6qdwYJru_Hgc7rRzFK9dJiCz3iDLCt10725HgrT_su9QF0DqFw7UravK-sanTjR6m4wcP2syBI1tFlDHht/s1034/middle%20east%20threat%201.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="628" data-original-width="1034" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgbH6i1tjAWCE0A1K3uud_J4NXP51n1xkPdikthYF6x171wUc7Fy5HHJZBRVZRfT-QwCdwR2IBR_U-BIRmWgB-hkehm5lpinAEC5XjVzm8Ktk6qdwYJru_Hgc7rRzFK9dJiCz3iDLCt10725HgrT_su9QF0DqFw7UravK-sanTjR6m4wcP2syBI1tFlDHht/w400-h243/middle%20east%20threat%201.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – A Summary of Where Most (But Not All) Middle
Eastern Violence Really Comes From<o:p></o:p></p>
<p class="MsoNormal"><u>It is my contention that Iran is ruled by a terroristic
dictatorship</u>, <u>masquerading as a
democracy</u>. The ruling figures pose
as religious mullahs to justify what they do,
but their actions clearly make that claim false. They can over-rule anything the “elected”
officials come up with, making it a sham
democracy. This group is propped up in
power by the Iranian Revolutionary Guard,
which is essentially their private army that ruthlessly suppresses
domestic dissent. <u>My contention
cannot be far from the truth</u>!<o:p></o:p></p>
<p class="MsoNormal">In point of fact,
there is very little operational difference between this situation, and 1933 Germany, with Hitler propped up by his private army, the brutal and ruthless SA. And we all know
where that led! I did generalize this to
a warning about leader cults, published
13 February 2020 as “Beware of Leader Cults” on this site. Such can be political, religious,
or both. To find any of these
quickly, use the navigation on the
left. Click the year, then the month, then the title if need be. <o:p></o:p></p>
<p class="MsoNormal">This terrorist government in Iran <u>funds</u>, <u>supplies</u>, and <u>commands</u> a bunch of proxy
terrorist armies in multiple places. Those
proxies are killing and wounding lots of people, and even killing and wounding US troops, as noted.
Iran already has cruise missiles and drones, as well as an army, a navy,
and an air force. It is going for
a nuclear weapon, and has been for some
time now. Iran is even launching
shipping attacks from within its own territory now, <u>so their vile behavior is escalating</u>. <o:p></o:p></p>
<p class="MsoNormal">Striking at proxies has proven ineffective, <u>because the
outfit giving the orders does not suffer any consequences</u>! This is indicated in Figure 2, where Russia has been unsuccessful for 2
years now swatting at the West’s proxy,
the Ukrainians. We in the West are
not generally terrorist dictatorships, not
yet, anyway. We should care a lot more about our proxy
than any dictatorship ever would, <u>and
therefore we all should see to it that they win</u>. Or overtly help them to win, if needed. <o:p></o:p></p>
<p class="MsoNormal">Economic sanctions against dictatorships have also proven
ineffective, because those do not hurt
the dictator, only those oppressed under
him, <u>and he cares not that they
suffer</u>! Our experiences with
Iran, Russia, North Korea,
and now China, all prove that
thesis. This is also indicated in the
figure.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_IAw-u58SvQcAqv7MhPyTVCsJF6kC7NnkRa-ze8jv4TOv1U5R-Dj4Xu5mCRjiDWNeJr4DjfdxrPpkgBRDQKsNN7eVSeoGibQWwclYTKmhgUMlg6ElM7PwFjmZoBDuKEiDX-9bUW8ynpHpCkBvizefu_94ciqFHjVBAGjSNz6Eeqi-Kj97T_HFID4yslGa/s1034/mideast%20action%202.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="628" data-original-width="1034" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_IAw-u58SvQcAqv7MhPyTVCsJF6kC7NnkRa-ze8jv4TOv1U5R-Dj4Xu5mCRjiDWNeJr4DjfdxrPpkgBRDQKsNN7eVSeoGibQWwclYTKmhgUMlg6ElM7PwFjmZoBDuKEiDX-9bUW8ynpHpCkBvizefu_94ciqFHjVBAGjSNz6Eeqi-Kj97T_HFID4yslGa/w400-h243/mideast%20action%202.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – What Does and Does Not Work, Dealing With Proxies, Plus a Suggestion<o:p></o:p></p>
<p class="MsoNormal"><span style="background: yellow; mso-highlight: yellow;">My firm
opinion: there will not be peace of any
kind in the Middle East until Iran is dealt with effectively. <b><i>Basically, that government must fall! </i></b> And none of us want to invade and occupy
Iran, we’ve already had enough of that
nonsense with Afghanistan and Iraq.</span><o:p></o:p></p>
<p class="MsoNormal">As the suggestion indicates,
one uses a surprise strike to kill as many of the ruling mullahs as
possible, in one fell swoop. After that,
one then conducts strikes from long range to destroy as many of the
Revolutionary Guard’s facilities and assets as can be found. <b><i>But you <u>do not invade</u>! </i></b>And leave the general infrastructure
alone. Just hit the Revolutionary Guard
stuff, and some of the military assets.<o:p></o:p></p>
<p class="MsoNormal">With the Revolutionary Guard thus weakened, and the ruling mullahs in disarray or
dead, there can be a successful popular
uprising, like the one about a decade or
so ago, that Mr. Obama failed to
support. The Iranian people are actually
good folk who would just as soon <u>not</u> be our enemies! This would give them a chance to be free and become
our friends (once again). <o:p></o:p></p>
<p class="MsoNormal"><b>Dealing with our multiple adversaries:<o:p></o:p></b></p>
<p class="MsoNormal">We the West face 4 really bad adversaries: Iran,
North Korea, Russia, and China.
Iran is now escalating from using all proxies into overt attacks, and they are close to having nuclear weapons. <b><i>They really do need to be dealt with
immediately</i></b>. <u>I just told you
how</u>.<o:p></o:p></p>
<p class="MsoNormal">North Korea is the one most likely to fling a nuke right
now, with Iran not far behind. The North Koreans don’t have very many
nuclear weapons yet, <u>so we can afford
some damage to take on Putin’s Russia next instead</u>, by forcing their utter defeat in, and complete expulsion from, Ukraine,
<u>any which way we can</u>! Putin
will be overthrown from within if that happens,
although there is no guarantee his successor won’t be just as evil as he
is. However, it will take time for the new one to
consolidate power, during which time we
can act.<o:p></o:p></p>
<p class="MsoNormal">Putin’s defeat and overthrow will tend to deter Xi’s
China, perhaps preventing World War 3 from
starting over Taiwan in the Pacific. His
country’s mounting economic stress (from going to a war production footing during
peacetime without much excuse other than starting a major war that the Chinese
people do not want) can possibly induce his overthrow from within, especially if that economy continues to falter
into serious recession. <o:p></o:p></p>
<p class="MsoNormal">With Russia and China out of the way, North Korea can be easily taken down. South Korea and Japan will be happy to
help. <u>That one IS an invasion</u>! It results in a reunified Korea, too.<o:p></o:p></p>
<p class="MsoNormal"><b>Final related remarks:<o:p></o:p></b></p>
<p class="MsoNormal">Do not be fooled by my inclusion of Israel among the nations
of the West. <span style="background: yellow; mso-highlight: yellow;">In the current war with Hamas, they have used tactics and weapons that
bespeak of a total disregard of Palestinian lives, well in excess above the appalling casualties
one would have to expect, from striking
at your enemy through their human shield.</span> It may actually qualify as some sort of war
crime. <o:p></o:p></p>
<p class="MsoNormal">That disregard of Palestinian lives is driven by the
far-right political coalition that currently governs Israel: they do not want a 2-state solution, they apparently want instead for there to be
no Palestinians left to have a state at all.
Very typical of authoritarian hard-liners! <b><i>The Israeli people need to change that
government, before it becomes a
dictatorship.</i></b> Netanyahu trying
to “reform” the Israeli judicial system was actually him trying to enable his staying
in power permanently. Israelis, you have been warned!<o:p></o:p></p>
<p class="MsoNormal"><span style="background: yellow; mso-highlight: yellow;">As for
the UN looking at war crimes charges against Israel: why are you not also looking at Hamas?</span> Using human shields is a war crime, as is attacking civilians! And Hezbollah, for dropping rockets on civilians for decades
now. What’s good for the goose is good
for the gander! <u>Hypocrisy at the UN stinks</u>!<o:p></o:p></p>
<p class="MsoNormal"><b><u><span style="background: yellow; mso-highlight: yellow;">As
for everybody else</span></u><span style="background: yellow; mso-highlight: yellow;">:</span></b><span style="background: yellow; mso-highlight: yellow;">
look out for your own right-wing extremists! They are among you!</span> They are invariably very authoritarian, and they invariably want to “eliminate”
(kill) their opposition. <b><i>If they
have committed any crimes, jail them for
it! Even if not, you need to vote them out, not in!</i></b> Or else you will have to revolt after they have
established their dictatorships over you.
Doing it at the ballot box instead,
leaves a whole lot less mess to clean up! <o:p></o:p></p>
<p class="MsoNormal"><span style="background: yellow; mso-highlight: yellow;">And
when I point my finger at you all, my
thumb is pointing back at me!</span> We
face the same peril here in the US, <u>with
multiple failed attempts (including an insurrection) to set up a right wing
extremist dictatorship</u>, <b><i>precisely
because we voted an authoritarian in,
instead of out, and he did not
want to give up power when his term in office was over. And now he wants another shot at it! Surprise,
surprise!</i></b><o:p></o:p></p><p class="MsoNormal"><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-1506923220063529672024-01-23T15:10:00.002-06:002024-01-31T09:22:44.544-06:00Trump Cult Warning<p>The following is the text of a signed letter I sent to PBS's NewsHour at their "viewermail" address. The letter is a warning about the threat of a coming dictatorship in the US. I am shouting from the rooftops about this threat! As for the image, some "peaceful tour or demonstration" that was!</p><p>------------------------------------</p><p>A much-edited version of this appeared as a column in the Waco, Texas, "Tribune-Herald" 30 January 2024</p><p>------------------------------------</p><p></p><p class="MsoNormal">I am still astounded that no one trying to responsibly
report the news seems to understand where all the rabidly-loyal Trump voters
are coming from,<span style="mso-spacerun: yes;"> </span>and why they persist in
their support, <span style="mso-spacerun: yes;"> </span>despite the serious criminal
trials of Trump that are coming.<span style="mso-spacerun: yes;"> </span>They so
very clearly want him, <span style="mso-spacerun: yes;"> </span>even if he is in
jail! <o:p></o:p></p>
<p class="MsoNormal">That’s because it is a <u>fearless leader cult</u> built
around Trump in the run-up to the 2016 election,<span style="mso-spacerun: yes;"> </span>from roots in the “Qanon” on-line
conspiracy.<span style="mso-spacerun: yes;"> </span>Cults are cults,<span style="mso-spacerun: yes;"> </span>belief outweighs the importance of any facts!<span style="mso-spacerun: yes;"> </span>They operate by brainwashing,<span style="mso-spacerun: yes;"> </span>and the Trump cult has been brainwashing GOP
voters since 2015.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Our cowardly GOP politicians know this,<span style="mso-spacerun: yes;"> </span>which is why they cravenly do whatever the
brainwashed crowd wants,<span style="mso-spacerun: yes;"> </span>no matter how
nonsensical or evil.<span style="mso-spacerun: yes;"> </span>They still outnumber
the true Trump cultists in the GOP,<span style="mso-spacerun: yes;"> </span>and
yet they are utterly subservient to them,<span style="mso-spacerun: yes;">
</span>because of those brainwashed voters. <span style="mso-spacerun: yes;"> </span>That is what “primarying” <u>really</u> means!<o:p></o:p></p>
<p class="MsoNormal">The GOP has morphed from a largely corporate-friendly
membership 3+ decades ago, <span style="mso-spacerun: yes;"> </span>to a major
percentage of uneducated blue-collar membership today,<span style="mso-spacerun: yes;"> </span><u>precisely because</u> uneducated people
are far easier to brainwash!<span style="mso-spacerun: yes;"> </span>The
brainwashing is done by both the politicians themselves at their rallies and
events,<span style="mso-spacerun: yes;"> </span>and by the “social media” on the
entirely unregulated internet,<span style="mso-spacerun: yes;"> </span>which
freely spouts all sorts of lies,<span style="mso-spacerun: yes;"> </span>and
even Russian (and other) propaganda!<span style="mso-spacerun: yes;"> </span>If
you don’t believe me,<span style="mso-spacerun: yes;"> </span>go look for
yourself!<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">That unregulated lying mess is where most GOP voters get
their “news” and their “facts”.<span style="mso-spacerun: yes;"> </span>If you
don’t believe me about that,<span style="mso-spacerun: yes;"> </span>go ask any
of them!<span style="mso-spacerun: yes;"> </span>But that unregulated lying mess
is also why they don’t believe any of the “mainstream media”,<span style="mso-spacerun: yes;"> </span>which is you,<span style="mso-spacerun: yes;">
</span>the responsible journalists still left in this world!<span style="mso-spacerun: yes;"> </span>And I think you do know about <u>that</u>!<o:p></o:p></p>
<p class="MsoNormal">Political cult,<span style="mso-spacerun: yes;">
</span>religious cult,<span style="mso-spacerun: yes;"> </span>or some of both,<span style="mso-spacerun: yes;"> </span>it does not matter.<span style="mso-spacerun: yes;"> </span>All have a purpose, <span style="mso-spacerun: yes;"> </span>and a top-of-the-list thing they want to
do.<span style="mso-spacerun: yes;"> </span>These are usually initially
secret,<span style="mso-spacerun: yes;"> </span>but they always come out.<span style="mso-spacerun: yes;"> </span>Trump’s former advisor,<span style="mso-spacerun: yes;"> </span>retired General Flynn, <span style="mso-spacerun: yes;"> </span>spilled the beans in public about the Trump
cult’s purpose,<span style="mso-spacerun: yes;"> </span>some years ago during
Trump’s term in the White House:<span style="mso-spacerun: yes;"> </span>a
dictatorship over the US,<span style="mso-spacerun: yes;"> </span>with Trump at
the top. <o:p></o:p><span style="color: red;">This was to start with the military
being ordered to confiscate voting machines,</span><span style="color: red;">
</span><span style="color: red;">then re-running the election to get the outcome the Trump cult
wanted.</span><span style="color: red;"> </span><span style="color: red;">(Updated in red 1-27-2024.)</span></p><p class="MsoNormal"><span style="color: red;"><o:p></o:p></span></p>
<p class="MsoNormal">Trump himself has made clear what his top-of-the-list
priority is, <span style="mso-spacerun: yes;"> </span>once he takes power: to
kill all his opposition,<span style="mso-spacerun: yes;"> </span>just not in
those exact words.<span style="mso-spacerun: yes;"> </span>His word is
“retribution”.<span style="mso-spacerun: yes;"> </span>But that’s what it
ultimately means!<span style="mso-spacerun: yes;"> </span>If you look at the
Qanon roots of the Trump cult,<span style="mso-spacerun: yes;"> </span>“the
storm” is where they round up all who oppose them (termed “deep state”),<span style="mso-spacerun: yes;"> </span>maybe (or maybe not) try them by a military
tribunal,<span style="mso-spacerun: yes;"> </span>not a real court,<span style="mso-spacerun: yes;"> </span>and execute them.<span style="mso-spacerun: yes;"> </span>That intent has not changed on the internet
since 2016;<span style="mso-spacerun: yes;"> </span>I have watched it.<o:p></o:p></p>
<p class="MsoNormal"><u>Liz Cheney was, <span style="mso-spacerun: yes;"> </span>and is, <span style="mso-spacerun: yes;"> </span>correct</u>:<span style="mso-spacerun: yes;">
</span><u>this nation is sleepwalking its way into a dictatorship</u>!<span style="mso-spacerun: yes;"> </span>The GOP has absolutely <u>no business</u>
running a multiply-indicted (and likely convicted) felon for the office of
President,<span style="mso-spacerun: yes;"> </span>or anything else!<span style="mso-spacerun: yes;"> </span>Especially one so evidently guilty of
fomenting an insurrection!<span style="mso-spacerun: yes;"> </span>That they are
running him is proof that the craven-coward politicians, <span style="mso-spacerun: yes;"> </span>and the Trump cult members of Congress, <span style="mso-spacerun: yes;"> </span>are together in total control of the Republican
party,<span style="mso-spacerun: yes;"> </span>which clearly is no longer a
responsible entity!<o:p></o:p></p>
<p class="MsoNormal">All the responsible media,<span style="mso-spacerun: yes;">
</span>NewsHour included,<span style="mso-spacerun: yes;"> </span>should be
screaming about this abomination from the rooftops!<span style="mso-spacerun: yes;"> </span>I certainly am.<o:p></o:p></p><p class="MsoNormal">------------</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEifBX520bFF9whgpp3fM11pZIUXbgn8Bn6lyOM4_1KbHxL_RXF2tRr7fihcQlIQ_Bip8aCx8zUz16vapnNhtcnEFsy7Zfp2ELHKVS9LqgyXVRkJBKqe6AcuUX-1KCCsSaT59s2egtRfUG26wY3yTfBFhk8QLMZ-9M76gnILP8ySP4I8pGSi6Pt0g0Ilo13x/s1600/trump%20insurrection.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1065" data-original-width="1600" height="266" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEifBX520bFF9whgpp3fM11pZIUXbgn8Bn6lyOM4_1KbHxL_RXF2tRr7fihcQlIQ_Bip8aCx8zUz16vapnNhtcnEFsy7Zfp2ELHKVS9LqgyXVRkJBKqe6AcuUX-1KCCsSaT59s2egtRfUG26wY3yTfBFhk8QLMZ-9M76gnILP8ySP4I8pGSi6Pt0g0Ilo13x/w400-h266/trump%20insurrection.png" width="400" /></a></div><p class="MsoNormal">------------</p><p class="MsoNormal">I have written articles before, on this site about the Trump cult. The best one is "Beware of Leader Cults", posted 16 February 2020. Use the navigation tool on the left to find it quickly. All you need is the date and the title. Click on the year, then the month, then the title if need be. </p><p class="MsoNormal">------------</p><p class="MsoNormal"><span style="color: red;">Update 1-27-2024: add another image as proof of roots. Look at caps, then at sign.</span></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjB960_fyh314RGhu633Pl2cgh3cOjbiMMRcAGPOqjXjEmGpTi0tX16rmvb06CpCZrZlNNlQ78BT7ggBJhA6w2BKhXuy7WEFCJPWYn0crUjfoR5SkwXCEvGWSmcjlTM2BEM1U6WT3TxVcHPap4_YDovYknC_YgvNu7h6U18ZqfMU-zfgWJ869REiDV-dguP/s806/MAGA%20as%20Q%202.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="465" data-original-width="806" height="231" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjB960_fyh314RGhu633Pl2cgh3cOjbiMMRcAGPOqjXjEmGpTi0tX16rmvb06CpCZrZlNNlQ78BT7ggBJhA6w2BKhXuy7WEFCJPWYn0crUjfoR5SkwXCEvGWSmcjlTM2BEM1U6WT3TxVcHPap4_YDovYknC_YgvNu7h6U18ZqfMU-zfgWJ869REiDV-dguP/w400-h231/MAGA%20as%20Q%202.png" width="400" /></a></div><br /><p class="MsoNormal"><br /></p><p class="MsoNormal"><br /></p><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-4014708944434202042024-01-08T11:51:00.001-06:002024-01-12T10:26:29.309-06:00Immigration Crisis?<p class="MsoNormal"><o:p></o:p></p><p>Yes, there is
one. No,
it is not the fault of Presidential administrations of either
party. It is primarily the fault of a
dysfunctional Congress not doing its sworn job.
See the figure.</p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjif-pl0BvTVb_xjcrk2W1Wcfx4Gw2Q2oSjHXSLPpOUVDr7KIF_6P6LcjdZDBw8gdFKLukOFsujRniHE2tZTA5S0-CIuzmIbv62Px3-g8jRaby3ISRlosHYIjm0x_MY_ugoOUEzr27ZsdtZbs_W_8rrYgEHdd_MPaglKjQRqP0yABoh8jZq1FjYxZK_BEKK/s1000/immigration.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="584" data-original-width="1000" height="234" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjif-pl0BvTVb_xjcrk2W1Wcfx4Gw2Q2oSjHXSLPpOUVDr7KIF_6P6LcjdZDBw8gdFKLukOFsujRniHE2tZTA5S0-CIuzmIbv62Px3-g8jRaby3ISRlosHYIjm0x_MY_ugoOUEzr27ZsdtZbs_W_8rrYgEHdd_MPaglKjQRqP0yABoh8jZq1FjYxZK_BEKK/w400-h234/immigration.png" width="400" /></a></div><p></p><p class="MsoNormal">This has been neglected since right after the end of World
War 2, except for one ineffective change
that George W. Bush got through Congress about 20 years ago. The public has become aware over the last 2
decades that there is a problem, but
Congress lies about whose fault it is,
instead of doing something constructive about it.<o:p></o:p></p><p class="MsoNormal">The guest worker quotas have always been set too low, since 1945.
Congress sets the maximums for the worker visa quotas. Administrations may issue less, but not more. Those visas cannot be tracked
effectively, because Congress never
appropriated the money to hire the people to do the tracking job. Why should anyone be surprised that, after decades of this mismatch, there are something like 10 million illegal residents in the US?<o:p></o:p></p><p class="MsoNormal">Until about 2 decades ago,
the number of guest workers we had to deal with was very much greater
than the number of asylum seekers we had to deal with. Under federal law, asylum seekers must be granted a hearing
before an immigration judge. The number
of these immigration courts and their infrastructure depends upon what Congress
funds. They were barely adequate 2+
decades ago. <o:p></o:p></p><p class="MsoNormal">The smugglers of drugs and people were only as numerous as
they were 2 decades ago, because of the
illegal guest workers to be smuggled.
Our border patrol was roughly adequate to the task then. No longer.<o:p></o:p></p><p class="MsoNormal">Since then, a large
number of failed states in South America,
the Caribbean, and elsewhere have
produced vast numbers of refugees seeking asylum, far more now, than we ever saw 20+ years ago! <b><i>The US has absolutely no control over
the conditions forcing those people to flee. </i></b> And under federal law, <u>they are guaranteed a hearing before an
immigration judge</u>; not entry, just the hearing. <u>That law says nothing about how they crossed the border to come apply</u>. <o:p></o:p></p><p class="MsoNormal">Administration policies requiring deportations without that
hearing are the Executive Branch literally violating federal law instead of
enforcing it. The wait time for an
immigration hearing now is a few to several years, because Congress has not funded an increase
in the number of immigration courts to address the vastly increased demand.<o:p></o:p></p><p class="MsoNormal">Walls and barriers demonstrably do not work nearly as well
as claimed. Even when they do work, their actual effect is to deny asylum seekers the
hearing they are entitled to, under our
laws. <b style="font-style: italic;">That denial is a violation of our
laws. </b>So are the stay-in-Mexico policies, and similar stuff.<o:p></o:p></p><p class="MsoNormal">What we need is a Congress that will update those laws and
fund the infrastructure that enables Presidential administrations to enforce
them faithfully. <o:p></o:p></p><p class="MsoNormal">What we have is a long history of dysfunctional Congresses blaming
administrations of the opposite party for the crisis, just to drum up votes. Playing politics is so very clearly prioritized way higher
than actually doing the jobs they swore to do.
It also shows up in Congress not having passed a proper budget in almost
20 years! Oath-breakers!<o:p></o:p></p><p class="MsoNormal">And bear in mind,
your own Congress has been lying egregiously about this immigration
crisis (and many other things) to you,
the citizens, for some decades
now. Both major parties do it, but one is currently worse than the other
about that. I’ll let you readers guess
which one. <o:p></o:p></p><p class="MsoNormal">Hint: the one that
wants the border walls everywhere.<o:p></o:p></p><p class="MsoNormal">There is a “fix” for this,
and it is coming up this very November! <b><i>Throw the bastards out! All
of them!</i></b> They certainly deserve it for
playing hardball politics instead of doing the jobs that they swore to do! Try somebody new, it does not matter who. <o:p></o:p></p><p class="MsoNormal">You cannot do any worse than what you have now. But you might do better.<o:p></o:p></p><p class="MsoNormal"><br /></p><p class="MsoNormal"><span style="background-color: #fcff01;"><b><u>Update 1-12-2024</u>:</b> It nauseates me to see the House GOP controlled by Trump-cult extremists who freely abuse any lever of power to get what they want, when they have not the votes, all the while lying so egregiously about what they are doing. My article above makes abundantly clear that the <b><i>immigration crisis is the fault of Congress's dereliction of duty over many decades</i></b>, and not the fault of the current President or any of his predecessors. Yet blaming Biden is the loudest noise being made by the Trump cult extremists dominating the House GOP. </span></p><p>
</p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com2tag:blogger.com,1999:blog-2675974463524895416.post-43227564305260706932024-01-02T20:58:00.000-06:002024-01-02T20:58:53.280-06:00Airplanes on Mars?<p class="MsoNormal">In a word, probably
not. <o:p></o:p></p>
<p class="MsoNormal">The mechanics required for steady flight are illustrated in <b>Figure
1</b> below. Basically, lift must balance weight, and thrust must balance drag. For clarity,
the moment balance about the center of gravity is not depicted. <o:p></o:p></p>
<p class="MsoNormal">Lift and drag are both proportional to the wind pressure and
the wing area. The coefficients of
proportionality are the lift and drag coefficients. Control of lift is by angle of attack
(AOA), and the usable lift coefficient
only varies between zero and the stall value (a bit over 1). For control purposes, the same basic lift coefficient values must
be used on Mars as on Earth.<o:p></o:p></p>
<p class="MsoNormal">The wind pressure (q, the dynamic pressure) is proportional
to density and to velocity <u>squared</u>.
The density can be calculated as a density ratio (σ)
multiplied by a standard density value (ρ<sub>o</sub>). Earth sea level density on a standard day is
the usual value used for that standard density.
Values are shown. <o:p></o:p></p>
<p class="MsoNormal">Under ideal gas assumptions (P = ρ R<sub>univ</sub> T/MW), the density ratio pretty much anywhere is the
pressure ratio to Earth standard pressure, multiplied by the molecular weight ratio to
standard, and divided by the <u>absolute-scale</u>
temperature ratio to standard. For
typical pressures and temperatures on Mars,
density ratio is near 1% of Earth sea level, but this is quite variable since the “air”
pressure there is quite variable. An
average value is shown. Surface gravity
on Mars is 38% that on Earth. <o:p></o:p></p>
<p class="MsoNormal">To design aircraft for Mars,
we need the surface density ratio divided by the surface gravity
ratio. The net effect is that the levels
of the aerodynamic forces acting on a reduced weight on Mars, are about factor
35 smaller than here on Earth, at
otherwise the same AOA’s! That factor
can increase either the wing area or the <u>square of the velocity</u>, or some of both, to get the same balance for steady flight on
Mars. <o:p></o:p></p>
<p class="MsoNormal">Note that if you make the wing bigger, it will be more massive in proportion to that
increased area, and therefore
heavier, even in the lower gravity of
Mars. The weight increase of more wing
area will act toward overcoming any lower speed benefit. That is because the density effect is much
larger than the reduced gravity effect,
on Mars.<o:p></o:p></p>
<p class="MsoNormal"><u>Velocity squared</u> factored up by 35 is the same as
velocity factored up by almost 6. Example:
if landing and takeoff speed for some airplane design was about 100 mph
on Earth, it would be almost 600 mph on
Mars for the same wing area as on Earth.
Such speeds that close to the surface are quite dangerous. That is just not something to be attempted
voluntarily. <o:p></o:p></p>
<p class="MsoNormal">Double the pressure to 12 mbar in the Hellas Basin, and the over-100 density reduction factor
halves. That density ratio divided by
the gee ratio is now closer to 17 than 35,
and its square root is a velocity ratio a bit over 4. It’s still a bigger, heavier (and impractical) wing by a factor of
17 on wing area, or else a speed near
the surface exceeding 400 mph. Or
something in-between, with an
impractically-large wing and a speed that is still too high close to the ground
to be safe. Not at all safe to attempt. Plus, you cannot fly it anywhere except down in
that basin.<o:p></o:p></p>
<p class="MsoNormal">The same basic aerodynamic and weight-carried factors act on
helicopter rotors in pretty much the same way.
<b><i>This is why I think the use of airplanes (or helicopters) as we
know them here on Earth, at a size scale
suitable for transporting freight or people,
are simply not technologically feasible in the extremely thin “air” of
Mars, despite the lower gravity. <o:p></o:p></i></b></p>
<p class="MsoNormal"><o:p></o:p></p><p>Not absolutely impossible,
but a practical design configuration is pretty much unimaginable. </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEijBzCLDyDtSU88_CiXUTOKJzZXrWxFOxReu0xopAW0pLjInkvK9-JSCNtH6-YKIRx9A532Zwms0W8jXxj86SnZL7CcAqSsOZBFtb2ZKbTJVT9eyDHsDfp8GhWo6UEo8h4TiLARz58_xQ7y5ddZggCeuzd-HOYyJVZjL9lcjmVTdGL-ZBC_DhVhNIbceJ8r/s1022/F1%20airplanes%20on%20Mars.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="614" data-original-width="1022" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEijBzCLDyDtSU88_CiXUTOKJzZXrWxFOxReu0xopAW0pLjInkvK9-JSCNtH6-YKIRx9A532Zwms0W8jXxj86SnZL7CcAqSsOZBFtb2ZKbTJVT9eyDHsDfp8GhWo6UEo8h4TiLARz58_xQ7y5ddZggCeuzd-HOYyJVZjL9lcjmVTdGL-ZBC_DhVhNIbceJ8r/w400-h240/F1%20airplanes%20on%20Mars.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – First Cut Exploration of Aircraft Design
Requirements for Mars<o:p></o:p></p>
<p class="MsoNormal"><b>Second, Closer
Look:<o:p></o:p></b></p>
<p class="MsoNormal">Now, looking at this
issue a bit more closely, let us explore
landing and takeoff speeds that are practical,
and at lift coefficients that are high,
but with adequate stall margin,
sort of like what is required by the FAR’s here on Earth. For that,
I presume 120 mph = 176 ft/sec = 53.6 m/s, and a max lift coefficient at takeoff and
landing of 1.0. I also looked at high
altitudes on Earth, and at higher
pressure in the Hellas Basin on Mars. I
did this with a spreadsheet, as
illustrated in <b>Figure 2</b> below. <o:p></o:p></p>
<p class="MsoNormal">Those numbers for Mars might not seem too bad, until you try to sketch what that kind of a change
to an aircraft design might look like. I
did that in <b>Figure 3</b> below,
holding the fuselage size constant,
and just up-sizing the wings and tails.
<u>There is no way to get the required tail arm lengths for stability
and control</u>, without also up-sizing
the fuselage, which drives up mass even
further! It would be the same with a
swept-wing design for higher cruise speeds:
<u>you still have to land and take off</u>! <o:p></o:p></p>
<p class="MsoNormal"><b><i>That should indicate just how impractical it will always
prove to be, to design conventional
airplanes capable of safe and practical flight on Mars, regardless of the propulsion. That “air” is just too thin!<o:p></o:p></i></b></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhy1O4FlR_-dASUG9NDWuHntJB0IPfM2R0Ce2SxyGOUCjRTKc5UsavAqB-Px9BpnGRTierOa1WRPkFclQKLCbnydIqQv-TFjvooRPFwxhzM5VntYO3qCEeFZUj76E0lZcX99MI6IR_vLUnEQN1Qq1cpiROq5hO1T5pIQYOsg0gnXvQLASqwpusysnINiHRt/s1047/F2%20wing%20for%20LTO.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="618" data-original-width="1047" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhy1O4FlR_-dASUG9NDWuHntJB0IPfM2R0Ce2SxyGOUCjRTKc5UsavAqB-Px9BpnGRTierOa1WRPkFclQKLCbnydIqQv-TFjvooRPFwxhzM5VntYO3qCEeFZUj76E0lZcX99MI6IR_vLUnEQN1Qq1cpiROq5hO1T5pIQYOsg0gnXvQLASqwpusysnINiHRt/w400-h236/F2%20wing%20for%20LTO.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Sizing Wings for 120 mph Takeoff/Landing Speeds
on Mars<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjL2m-t1m_C09HYO9V4Ru7ewDs69YrIo-jIj1bfxTeuMIW-5PAeqWTlc1GAgQGEJz2FbW6Vz-t_5h0VOAn4eSa2RAq0kDV_YQS3W-Bd5P9Zv4p2OD5buMNuOQtLEY2XxoU7eJGO4gAbEtnVTwmjE-K6OAl3YPTvy7XAJw5sEsq_FZGgmxKSa4-EozBVFx8A/s1047/F3%20upsize%20aerosurfaces.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="618" data-original-width="1047" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjL2m-t1m_C09HYO9V4Ru7ewDs69YrIo-jIj1bfxTeuMIW-5PAeqWTlc1GAgQGEJz2FbW6Vz-t_5h0VOAn4eSa2RAq0kDV_YQS3W-Bd5P9Zv4p2OD5buMNuOQtLEY2XxoU7eJGO4gAbEtnVTwmjE-K6OAl3YPTvy7XAJw5sEsq_FZGgmxKSa4-EozBVFx8A/w400-h236/F3%20upsize%20aerosurfaces.png" width="400" /></a></div><p class="MsoNormal">Figure 3 -- Upsizing
Aerosurfaces for Fixed Fuselages For Mars<o:p></o:p></p><p><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-12782097233434692432023-12-29T17:12:00.002-06:002024-01-12T10:20:22.296-06:00Israel vs Hamas: It Is Worse Than You Think<p>It does not matter whether you are a Palestinian supporter
or an Israeli supporter, you are likely
wrong in your beliefs, precisely because
you do not understand how complicated all of this really is. There’s more than enough blame to go around
for both sides, and yet more blame besides
that.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">The <b>figure below</b> shows a map of the Palestine region
as it was proposed to be divided between Jews and Palestinians by the UN’s 1947
partition proposal.<span style="mso-spacerun: yes;"> </span>That was a 56-44
split of area between Jews and Palestinian Arabs,<span style="mso-spacerun: yes;"> </span>with Jerusalem intended to be an open city
belonging to neither side. While not an even split,<span style="mso-spacerun: yes;"> </span>the Jewish area included a lot of the
more-or-less uninhabitable Negev desert in the south.<o:p></o:p></p>
<p class="MsoNormal">The neighboring Arab nations did not agree to this,<span style="mso-spacerun: yes;"> </span>but it was pretty well done anyway by the UN
and western nations in 1948, <span style="mso-spacerun: yes;"> </span>with the
creation of Israel,<span style="mso-spacerun: yes;"> </span>pretty much along
the proposed lines pictured.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Almost immediately,<span style="mso-spacerun: yes;">
</span>the neighboring Arab nations invaded,<span style="mso-spacerun: yes;">
</span>entering through the Palestinian territories,<span style="mso-spacerun: yes;"> </span>starting the 1948 war,<span style="mso-spacerun: yes;"> </span>which they ultimately lost,<span style="mso-spacerun: yes;"> </span>losing some of the Palestinian territory to
the Israelis,<span style="mso-spacerun: yes;"> </span>as indicated in the
figure,<span style="mso-spacerun: yes;"> </span>but keeping the bulk of it as
shown.<span style="mso-spacerun: yes;"> </span><b><i>Although they could
have,<span style="mso-spacerun: yes;"> </span>these neighboring Arab states <u>never</u>
gave the Palestinian territories they occupied back to the Palestinians!<o:p></o:p></i></b></p><p class="MsoNormal"><b></b></p><div class="separator" style="clear: both; text-align: center;"><b><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhNviZCgdleXs6kx6Mq51NCN-7ZTc8WgQeujdMwdySu-U0NIeRpWSHJZIm8WpApGO89xAVgYR30dDChFWI_lE4wr8jfxd078UpsIYPdY3mGWpBlwHHsm_e_xbikMLhhAcFbj462SwscJxGm8CnTErJLdWZjpKls3lp4IrTYPNv0iys9fHysBvBljCLw_xhp/s1241/320px-1947-UN-Partition-Plan-1949-Armistice-Comparison.svg.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="898" data-original-width="1241" height="290" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhNviZCgdleXs6kx6Mq51NCN-7ZTc8WgQeujdMwdySu-U0NIeRpWSHJZIm8WpApGO89xAVgYR30dDChFWI_lE4wr8jfxd078UpsIYPdY3mGWpBlwHHsm_e_xbikMLhhAcFbj462SwscJxGm8CnTErJLdWZjpKls3lp4IrTYPNv0iys9fHysBvBljCLw_xhp/w400-h290/320px-1947-UN-Partition-Plan-1949-Armistice-Comparison.svg.png" width="400" /></a></b></div><p></p><p class="MsoNormal"><b>Multiple Wars Since<o:p></o:p></b></p><p class="MsoNormal">There were wars in 1951-1955 (the Palestinian Fedayeen
insurgency) and in 1956 (with Egypt over the Suez Canal) that did not
materially affect the map shown in the figure.
It was essentially unchanged until the 1967 “6-Day War”, when Israel captured the Golan heights in the
north from Syria, the Gaza Strip and
Sinai Peninsula from Egypt, and much of
the West Bank from Jordan in the east.
The Sinai Peninsula was eventually returned to Egypt as part of the
peace accords after the 1973 “Yom Kippur” war.
<o:p></o:p></p><p class="MsoNormal">1967-1970 saw the so-called “War of Attrition”, and 1971-1983 saw the Palestinian insurgency
in south Lebanon. The map didn’t really
change, with the Gaza Strip and much of
the West Bank occupied. 1973 saw the
“Yom Kippur” war, resulting in peace
accords and recognition of Israel by Egypt.
<o:p></o:p></p><p class="MsoNormal">Since then, there was
the 1978 First South Lebanon conflict,
resulting in the Palestine Liberation Organization (PLO) being expelled
from <u>southern</u> Lebanon. (That’s
the same PLO that is today the government of the West Bank.)<o:p></o:p></p><p class="MsoNormal">There was the 1982 First Lebanon War, resulting in the PLO being expelled from <u>all</u>
of Lebanon. There was also the 1985-2000
South Lebanon conflict, which resulted
in Hezbollah being based permanently in Lebanon (and Syria). <o:p></o:p></p><p class="MsoNormal">Then there was the 1987-1993 First Palestinian
Intifada, which the Israelis
suppressed. During it were the 1991
Iraqi rocket attacks on Israel, which
was a strategic failure for Iraq, since
it did not provoke a response. 2000-2004
was the Al Aqsa (or second) Intifada,
also suppressed. <o:p></o:p></p><p class="MsoNormal">Then in 2006 was the Second Lebanon War, where the Israelis took on both Hezbollah and
the Lebanese military fighting together.
That was inconclusive, but with
large combat losses.<o:p></o:p></p><p class="MsoNormal">All of that is <u>part</u> of the background to the current
2023 Israeli-Hamas war that ongoing,
with war also starting in the north with Hezbollah in Lebanon.<o:p></o:p></p><p class="MsoNormal"><b>Related Stuff<o:p></o:p></b></p><p class="MsoNormal">Another part is the history of Israeli settlements in the
West Bank. While the territory was
occupied, it was <u>never officially
annexed</u>, because if it were, the Palestinian Arabs there would add to
those living within Israel as Israeli citizens,
and outnumber the Jewish citizens.
Not annexed, those settlements
may well be technically illegal,
depending upon your interpretations of international law and UN resolutions.<o:p></o:p></p><p class="MsoNormal">Yet another part of this is the far-right nature of the
coalition that is Netanyahu’s government in Israel. It was that way when he was Prime Minister before, and it is that way now that he is again. When he was Prime Minister before, that’s when the pace of planting settlements
in the West Bank vastly accelerated.
Those settlers would by-and-large be of a similar far-right mindset, just to choose to settle there, the whole point being to effectively make
that Israeli land, no matter what.<o:p></o:p></p><p class="MsoNormal">He is Prime Minister again,
and so effectively in command of the Israeli military. The “court reform” is him trying to pave the
way to staying in ever-more-power,
permanently. The settlers in the
West Bank that he essentially put there,
are now going around killing their Palestinian sheep-herding neighbors
in this latest war. And the Israeli
military in Gaza is using strategy and tactics that reflect an utter disregard
for Palestinian civilian lives. Armies
often tend to reflect their commanders-in-chief. <o:p></o:p></p><p class="MsoNormal"><b><i>Why would anyone be surprised at that behavior, given who and what Netanyahu is?<o:p></o:p></i></b></p><p class="MsoNormal">And, <b><i>no one
should be surprised that the Arab neighbor nations are refusing to take in
Palestinian refugees from the Gaza Strip</i></b>, given the fact they never gave back the lands
in 1948. That border crossing with Egypt
is closed. It has been difficult to get
them to allow UN aid trucks in. <o:p></o:p></p><p class="MsoNormal"><u>What is surprising</u> is that no one is reporting on all
this ugly history and background! <o:p></o:p></p><p class="MsoNormal">And don’t forget what Hamas did with its October 7<sup>th</sup>
attack of unprecedented scope and evil. They
took hostages, killed with great
barbarity including women and children,
and raped a lot of women. <b><i>Those
are internationally-recognized war crimes,
as is hiding behind human shields.</i></b> <o:p></o:p></p><p class="MsoNormal">You should also be unsurprised at the civilian casualties that
inevitably happen when you strike back at an enemy like that. You have to go through that human shield to
strike them. It is inherent. And, <b><i>they tried to prevent their own civilians
from evacuating</i></b>, preferring
instead that they die serving as human shields.
<o:p></o:p></p><p class="MsoNormal">And THAT evil is the government of Gaza? They do not care whether the people they
govern live or die!<o:p></o:p></p><p class="MsoNormal"><b style="background-color: #fcff01;">My Predictions<o:p></o:p></b></p><p class="MsoNormal"><span style="background-color: #fcff01;">Hamas will fight until it is destroyed, but at high losses to the Israeli
military, and utterly enormous losses of
civilian Gazan lives. Hezbollah is lots
bigger and better equipped than Hamas,
and will continue the war with Israel in the north, after Israel’s forces have been damaged by
Hamas.<o:p></o:p></span></p><p class="MsoNormal"><span style="background-color: #fcff01; color: red;"><b><u>Update 1-12-2024</u>: </b> these two predictions have already come true. Which makes the following prediction even more crucial:</span></p><p class="MsoNormal"><span style="background-color: #fcff01;">Both Hamas and Hezbollah are well-known to be terrorist
proxy armies funded, supplied, and commanded by the government of Iran, <u>which has so far suffered precisely zero
consequences for what its proxies have been doing</u>. This violence is <u>not going to stop</u>
until they do suffer for what they have done.
</span><o:p></o:p></p><p class="MsoNormal"><b>List of Iranian Terrorist Proxy Armies <o:p></o:p></b></p><p class="MsoNormal">Iran is a terrorist dictatorship run by fake mullahs misusing
religion to “justify” what they do. They
run a sham democracy in which they can over-rule anything an elected body
decides, they suppress dissent
violently, and they are essentially
propped-up in power by their private army the Revolutionary Guard. Sounds an awful lot like Hitler and his SA stormtrooper
private army in 1930’s Germany to me!<o:p></o:p></p><p class="MsoNormal">Iran also operates <u>multiple terrorist proxy armies.</u><o:p></o:p></p><p class="MsoNormal" style="margin-bottom: 0in;">Per “Country Reports on Terrorism:
Iran 2021” by Bureau of Counterterrorism at US Dept. of State,<o:p></o:p></p><p class="MsoNormal">retrieved from their site 19 November 2023 (highlighting is
mine, for this article):<o:p></o:p></p><p class="MsoNormal">
<b></b></p><div class="separator" style="clear: both; text-align: center;"><b><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgLrgSH99zaxkGIaDcy-9Um6ZJ_NTnglp7rpN7-A-raqdSZnp56OJ-mYo4wF0e9Br1Rng-fSjcYBIXu-5MNAe6GyMyc62-1HncQgt5YwyJsoBD1_fBapQQxo2V6TeQzSktaoikN6F4MsIcRUhE56Pzto28EkH01kKbch1VivGG6zKkeusVpYJbD_A3G5foD/s1010/Iranian%20proxies.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="274" data-original-width="1010" height="109" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgLrgSH99zaxkGIaDcy-9Um6ZJ_NTnglp7rpN7-A-raqdSZnp56OJ-mYo4wF0e9Br1Rng-fSjcYBIXu-5MNAe6GyMyc62-1HncQgt5YwyJsoBD1_fBapQQxo2V6TeQzSktaoikN6F4MsIcRUhE56Pzto28EkH01kKbch1VivGG6zKkeusVpYJbD_A3G5foD/w400-h109/Iranian%20proxies.png" width="400" /></a></b></div><p></p><p class="MsoNormal">I would also point out the <u>recent attacks on commercial
shipping</u> in that same region. Many
of these are being conducted from Yemen by the Houthis, and recently,
there was a direct attack from Iran itself by the Revolutionary
Guard, at a ship off the coast of India!
<o:p></o:p></p><p class="MsoNormal"><b style="background-color: #fcff01;">Wake Up and Do This Right<o:p></o:p></b></p><p class="MsoNormal"><span style="background-color: #fcff01;">This violence will never end, as long as Israel or anybody else just swats
at these terrorist proxy armies. It can
only end when the terrorist government of Iran is overthrown. I have nothing at all against the people of
Iran, they are good people. <u>But their terrorist government must go</u>! Or there can NEVER be peace in the Missile
East!</span><o:p></o:p></p><p class="MsoNormal"><b>You Will Not Like Hearing This, Either:<o:p></o:p></b></p><p class="MsoNormal"><u>Ukraine is the West’s proxy against the Russians</u>. That has dragged-on nearly 2 years. It is time to do more than just send aid and
money, Russia has to lose, and soon!
Which will likely cause regime change in Russia. And that would be a good thing, even if it did not affect things with China!<o:p></o:p></p><p class="MsoNormal">But this DOES affect China!
If Russia does not lose, and
soon, Xi in China will be further
emboldened by the West’s apparent weakness,
and will start World War 3 in the Pacific by invading Taiwan. <o:p></o:p></p><p class="MsoNormal"><b><i><span style="background-color: #fcff01;">And the longer this Ukraine thing drags on, the more likely he will invade Taiwan
anyway, whether Russia wins or loses.</span><o:p></o:p></i></b></p><p class="MsoNormal">So just wake up and get on with it! <o:p></o:p></p><p class="MsoNormal">Proxy army war can get complicated, can it not?<o:p></o:p></p><p class="MsoNormal"><o:p> </o:p></p><p class="MsoNormal">
<b><i><br /></i></b></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-64941970484090438482023-12-16T18:54:00.003-06:002023-12-18T09:37:25.096-06:00Christmas Funnies 2023 <p>This is just for fun.
My wife found a funny on the internet that I simply must repost
here. The lights as hung in this city
display are just too funny, given the
appropriate modified Christmas song to go with them. </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgFH2NxWHnv3cPZodw3YgFO9Ome7eeZE9egcJJ2kVAcbgHwqgbJ74Fej7X1BcXQDFXilDtDRWjxk3LFgDpT8o2LMlGmuZVlpace5V5_79GS7uz90MfnTrvGUlGbsBhS5tlt-cC78WhjQTGjrlXR6cE919sn7Bb76uDvIjm1Aa_bKvfpBBlMsGYIwC2OYRfs/s574/xmas%20funny.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="574" data-original-width="526" height="400" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgFH2NxWHnv3cPZodw3YgFO9Ome7eeZE9egcJJ2kVAcbgHwqgbJ74Fej7X1BcXQDFXilDtDRWjxk3LFgDpT8o2LMlGmuZVlpace5V5_79GS7uz90MfnTrvGUlGbsBhS5tlt-cC78WhjQTGjrlXR6cE919sn7Bb76uDvIjm1Aa_bKvfpBBlMsGYIwC2OYRfs/w366-h400/xmas%20funny.png" width="366" /></a></div><p class="MsoNormal">As many readers know,
I put up a yard display each year that we call “White Trash
Christmas”. It involves some plastic
yard flamingoes lit up from inside, as
Santa’s reindeer, a lit-up plastic Santa, and something I call the “Iron Christmas
Tree”. There are two views of this
posted below, taken just after sun-down, Saturday 12-16-2023. I put a trash bag of recycled cans in the
wagon to represent Santa’s magic bag. <o:p></o:p></p>
<p class="MsoNormal">There are 10 flamingoes.
Eight are arranged in pairs,
marked and ordered as the eight names from the Christmas tales. Those would be “Dasher”, Dancer”,
Prancer”, and ”Vixen”, followed by “Comet”, “Cupid”, “Donner”,
and “Blitzen”. The lone one out
front is labeled “Rudolph”. The lone one
behind is labeled “Bambi”. Yeah, I cheated.
But I had 10 yard flamingoes. <o:p></o:p></p>
<p class="MsoNormal">The Iron Tree is something I made many years ago. It has a PVC pipe center core, and angled “legs” made of rebar. The lights get hoisted up and then
“may-poled” around, to cover it. <o:p></o:p></p>
<p class="MsoNormal">The view from more-or-less the side shows this thing set up
near my farm shop. The view from behind
shows our house across the driveway from the farm shop. <o:p></o:p></p>
<p class="MsoNormal">By the way, each
flamingo has pipe cleaner “antlers” on his head. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj5357Ggd8AbH5qN7c1DAP9zo3aKzRa-7MseKNhw9S4-cwCHmP47RnDfO2jvM680bNX9vMQqhdG8s8Lr-SQBmsk1ux9BnkuyrHtGrwCI2_vx6oMZaLdSMnSexxu4HGHduk1fpft-xFWwdBgZJaMFxbosz55tQOIL5X5r1ET4bS7x5SNrmZomYF9NXtuOo6h/s1600/IMG_0892.jpg" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1600" data-original-width="1200" height="640" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj5357Ggd8AbH5qN7c1DAP9zo3aKzRa-7MseKNhw9S4-cwCHmP47RnDfO2jvM680bNX9vMQqhdG8s8Lr-SQBmsk1ux9BnkuyrHtGrwCI2_vx6oMZaLdSMnSexxu4HGHduk1fpft-xFWwdBgZJaMFxbosz55tQOIL5X5r1ET4bS7x5SNrmZomYF9NXtuOo6h/w480-h640/IMG_0892.jpg" width="480" /></a></div><br /><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiesSNLR86hGeihlmQGd0GwJQTzbPjCAvj4H59rHR1Fsn8gYLgWMLGHL7r3lgAT6IQ2cU8QbtriiUlssWs4KsJfAOCaERLJxJPAmS77ZPnvvYXyCBfc8A2RIhGdg7LwXbMFUjGVi5QwWZf0Y8LqQAQ4X2YRnPxPuUc51MYm7YxO0QhHS00ulzjRIoz0wCUx/s1600/IMG_0894.jpg" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1600" data-original-width="1200" height="640" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiesSNLR86hGeihlmQGd0GwJQTzbPjCAvj4H59rHR1Fsn8gYLgWMLGHL7r3lgAT6IQ2cU8QbtriiUlssWs4KsJfAOCaERLJxJPAmS77ZPnvvYXyCBfc8A2RIhGdg7LwXbMFUjGVi5QwWZf0Y8LqQAQ4X2YRnPxPuUc51MYm7YxO0QhHS00ulzjRIoz0wCUx/w480-h640/IMG_0894.jpg" width="480" /></a></div><div><br /></div><b><u>Update 12-18-2023</u>:</b> We also put up this year's version of the missile toad. <div><br /></div><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEinNgxgZ0CYRPa0lLRQ3VdDVqFNSL71VElpHyBUftrqT5ZW8D1x60UiNHh8Xr4QNvCFxIWqU6P9Wjue0iiv8B4Z0qvaa8peiorfqMHCzq-rxgMjoKhQwBxeHM-1mpv2_3LaTHaNjVrQqlyIc3aKEDgD9HE-CIuk075RS5tpRwaVS9pTswgWd_Il5R9zET3M/s1600/IMG_0896.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1600" data-original-width="1200" height="640" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEinNgxgZ0CYRPa0lLRQ3VdDVqFNSL71VElpHyBUftrqT5ZW8D1x60UiNHh8Xr4QNvCFxIWqU6P9Wjue0iiv8B4Z0qvaa8peiorfqMHCzq-rxgMjoKhQwBxeHM-1mpv2_3LaTHaNjVrQqlyIc3aKEDgD9HE-CIuk075RS5tpRwaVS9pTswgWd_Il5R9zET3M/w480-h640/IMG_0896.jpg" width="480" /></a></div>And just to let you know that not everything around here is abnormal, our tree is up, too.<br /><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjsD2yJz2cU1UXgkH5otmOmhKa8wHqTbQF5EdZM65YqpHl-JCrFx8GNv86DR1DJff1HeSidNpGXAoPEP2lb2coabFvgaL7LGHOy01h50VZBQP0DQxg-QZkYwiwOD38PKLc40trRa-OXrNVIBIEmSN8md1KXQYOHYIvwWhlKt_6R5zTPNX-ib8g_m5WJM0jZ/s1600/IMG_0897.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1600" data-original-width="1200" height="640" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjsD2yJz2cU1UXgkH5otmOmhKa8wHqTbQF5EdZM65YqpHl-JCrFx8GNv86DR1DJff1HeSidNpGXAoPEP2lb2coabFvgaL7LGHOy01h50VZBQP0DQxg-QZkYwiwOD38PKLc40trRa-OXrNVIBIEmSN8md1KXQYOHYIvwWhlKt_6R5zTPNX-ib8g_m5WJM0jZ/w480-h640/IMG_0897.jpg" width="480" /></a></div><br /><div><br /></div><div><br /></div><div> </div><div><br /><p><br /></p><p class="MsoNormal"><o:p></o:p></p></div>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-70443717386255397032023-12-10T15:47:00.000-06:002023-12-10T15:47:06.006-06:00Conspiracies Have Short Lifetimes<p>As most readers of this blog already know, I have long denounced the conspiracy theory
that there is some evil “deep state” in control of our government and others
around the world. This derives largely
from Qanon and similar or related sites on the internet, all very popular with right-wing and
far-right readers. While some aspects
and pieces of this movement date earlier,
this became a fearless leader cult around Donald Trump during 2015, as he began his run for the Presidency in the
2016 election. </p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">Key to this is supposedly secret information about forces
led by Trump to depose this “deep state”.<span style="mso-spacerun: yes;">
</span>This is to take place as a “storm” during which “deep state” people will
be rounded up,<span style="mso-spacerun: yes;"> </span>imprisoned,<span style="mso-spacerun: yes;"> </span>perhaps tried or perhaps not,<span style="mso-spacerun: yes;"> </span>but definitely executed.<span style="mso-spacerun: yes;"> </span>Invariably,<span style="mso-spacerun: yes;">
</span>the persons identified as targets of this “storm” are Hollywood elites,<span style="mso-spacerun: yes;"> </span>Democrats all over,<span style="mso-spacerun: yes;"> </span>and any Republicans who are not Trump
cultists (referred to as RINO’s). <span style="mso-spacerun: yes;"> </span>Also
invariably,<span style="mso-spacerun: yes;"> </span>cultists refer especially to
Democrats as being evil enemies of the state,<span style="mso-spacerun: yes;">
</span>worthy only of being killed,<span style="mso-spacerun: yes;">
</span>which has inspired considerable “lone wolf” violence of late. <span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">To accomplish this requires overthrowing our democracy and
replacing it with a Trump-led dictatorship.<span style="mso-spacerun: yes;">
</span>That’s the secret!<span style="mso-spacerun: yes;"> </span>Which is
exactly why so many Trump cultists have become publicly known to be “authoritarians”
(meaning advocates of a dictatorship in America).<span style="mso-spacerun: yes;"> </span>It is why Trump himself has recently begun saying
“dictatorship on only day one” if re-elected.<span style="mso-spacerun: yes;">
</span>The only part of that which is a lie,<span style="mso-spacerun: yes;">
</span>is the word “only”.<span style="mso-spacerun: yes;"> </span>And given his
track record over the decades for egregious lying,<span style="mso-spacerun: yes;"> </span>that is no surprise at all!<o:p></o:p></p>
<p class="MsoNormal">Secrets are hard to keep,<span style="mso-spacerun: yes;">
</span>and with conspiracies to keep them secret,<span style="mso-spacerun: yes;"> </span>there are invariably some numbers of people
who know about them.<span style="mso-spacerun: yes;"> </span>It has always been
the pattern in the past,<span style="mso-spacerun: yes;"> </span>that eventually
somebody talks and spills the secret.<span style="mso-spacerun: yes;">
</span>Thus there is some sort of finite “lifetime” that such information,<span style="mso-spacerun: yes;"> </span>and the conspiracy to keep it secret, <span style="mso-spacerun: yes;"> </span>both have.<o:p></o:p></p>
<p class="MsoNormal">I ran across a news item on the PBS NewsHour web site that
describes the modeling of such conspiracy lifetimes.<span style="mso-spacerun: yes;"> </span>This was “How Many People Does It Take to
Keep a Conspiracy Alive?”,<span style="mso-spacerun: yes;"> </span>which I
retrieved from their web site 7 December 2023.<span style="mso-spacerun: yes;">
</span>It was apparently published there 15 February 2016,<span style="mso-spacerun: yes;"> </span>but I missed it back then.<span style="mso-spacerun: yes;"> </span>The article reports the published academic
work of one David Robert Grimes at Oxford University,<span style="mso-spacerun: yes;"> </span>who found an equation modeling this effect as
lifetime-to-revelation versus number-of-people-involved.<span style="mso-spacerun: yes;"> </span>He reportedly published this equation in the
Journal PLOS ONE. <o:p></o:p></p>
<p class="MsoNormal">I did not track down his published article.<span style="mso-spacerun: yes;"> </span>The PBS NewsHour article gave enough
numerical data for me to just reverse-engineer an approximation to what Grimes
apparently found and published.<span style="mso-spacerun: yes;"> </span>I simply
keyed the quoted data into a spreadsheet,<span style="mso-spacerun: yes;">
</span>plotted and re-plotted the data multiple ways,<span style="mso-spacerun: yes;"> </span>and attempted some curve-fit equations.<span style="mso-spacerun: yes;"> </span>One of those worked rather well,<span style="mso-spacerun: yes;"> </span>and is probably fairly close to the Grimes
result.<span style="mso-spacerun: yes;"> </span>I even found one “bad data point”
item quoted in the PBS news article.<span style="mso-spacerun: yes;"> </span>My
work is depicted by the plots in Figure 1. <span style="mso-spacerun: yes;"> </span>It is the log-log plot options that produced
fittable curves. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgTbDw3Ts6jukQKMbZQrd-_VAyZ6B2EaU6MhbQdjbctoQegAdN2wSmYzwjmQT_SqAXLu0UO6hQaHPUDqIRqJnYkH1N6lS9x_JhMt3xaOvOf-bipzPKBY-WK2XqEZVXU1pak3w0R6ow0Ys1xlySdr0YONeaseBE4g15PewWbgKpFa0FCLih9vLohpCa0ctEj/s1031/consp%20life%20eqn.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="624" data-original-width="1031" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgTbDw3Ts6jukQKMbZQrd-_VAyZ6B2EaU6MhbQdjbctoQegAdN2wSmYzwjmQT_SqAXLu0UO6hQaHPUDqIRqJnYkH1N6lS9x_JhMt3xaOvOf-bipzPKBY-WK2XqEZVXU1pak3w0R6ow0Ys1xlySdr0YONeaseBE4g15PewWbgKpFa0FCLih9vLohpCa0ctEj/w400-h243/consp%20life%20eqn.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Reverse-Engineering a Correlation Equation From
Quoted Data (Excluding “Bad Data Point”)<o:p></o:p></p>
<p class="MsoNormal">The modeling equation that worked best, as indicated by the higher Pearson’s
r-squared parameter, is in the upper
right plot of the figure. It models the
data quite well. It is:<o:p></o:p></p>
<p class="MsoNormal"> Y =
4.4378 [log<sub>10</sub>(no. of ppl)]<sup>-1.235</sup><o:p></o:p></p>
<p class="MsoNormal"> Life,
yrs = 10<sup>Y</sup><o:p></o:p></p>
<p class="MsoNormal">Now, apply this to
the Trump-cult (Qanon) conspiracy, whose
“secret” is that they want a Trump-led dictatorship in the US, and also that they want to kill all their
opposition! This formed about 8 years
ago in 2015 during the campaign leading to the 2016 election. Solving iteratively the equation-in-reverse
for a 8-year life, then no more than
about 4262 persons could know about the secret,
and still have it remain hidden the 8 years since then. <u>That’s clearly not the case</u>, as there are millions of Trump supporters who
voted for him in both 2016 and 2020. At
least some of them are cult believers.<o:p></o:p></p>
<p class="MsoNormal">Alternatively, in
those elections, there were 80-something
millions of Trump voters. I picked 82
million as an estimate, just to be the
right magnitude only. <u>Assuming all of
them are part of the Trump-cult conspiracy</u>,
the equation says the expected lifetime before the secret gets revealed
would be only about 2.2 years. The
assumption I made is not correct, but
using it helps bound the problem between these two results: 2.2 vs 8 years after the 2015 campaign. The “right” answer must lie between these
points.<o:p></o:p></p>
<p class="MsoNormal">In actuality, the
revelation that Trump and his advisors were integrally associated with the
Qanon theory came out during his Presidential administration January
2017-January 2021. Trump’s then-adviser
retired General Flynn <u>publicly advocated</u> for Trump to replace our
democracy with a dictatorship led by Trump,
so that the “deep state” could be “overthrown” and eliminated! This really means all their opposition being rounded
up and killed, exactly like what
happened when Adolf Hitler’s Nazis took over Germany. <u>This revelation happened about 4 years
ago, exactly between the bounds I just
defined</u>!<o:p></o:p></p>
<p class="MsoNormal">Using the equation iteratively-in-reverse again, on the 4 years since the 2015 campaign to
sometime in 2019, it estimates the
number of really hard-core, extremist
Trump-cult believers to be about 109,700.
While not millions, that’s large
enough to be very alarming, as the pool
from which right-wing extremist violence can most likely be expected to
come. And we have already seen it.<o:p></o:p></p>
<p class="MsoNormal">And, we already have several
Representatives and a few Senators who so very clearly come from this
cult. So, why is our Congressional dysfunction a
surprise? It’s part of their plan! Chaos instead of governing helps to “justify”
their radical change to a dictatorship under Trump.<o:p></o:p></p>
<p class="MsoNormal">Subtracting that roughly 110,000 true-believer cultists from
about 82 million evident Trump voters,
that says <b><i>something like 81.9 million voters have been deceived by
this cult, into voting for a would-be
dictator over the US, who has already
attempted an insurrection and takeover on 6 January 2021</i></b>. <o:p></o:p></p>
<p class="MsoNormal"><b><u><span style="background: yellow; mso-highlight: yellow;">And
THAT ought to scare the ever-loving shit out of almost anybody</span></u><span style="background: yellow; mso-highlight: yellow;">!</span><o:p></o:p></b></p>
<p class="MsoNormal">Please wake up out there!<o:p></o:p></p><p class="MsoNormal"><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-3586956950947121022023-12-09T18:17:00.001-06:002023-12-10T17:06:57.564-06:00Overall Study Results: Propellant From Moon<p class="MsoNormal"><o:p></o:p></p><p>Just the overall results are given here. <u>The details supporting this are much more
voluminous</u>. The basic notion, Figure 1,
is to manufacture propellants on the moon, probably using potentially-recoverable ice
deposits near the south pole. I
initially looked at elongated halo orbits about the moon as a means to more
easily access the south polar region off the moon from orbit. Delivery is to low Earth orbit, at low inclination eastward. The idea is to base a lander on the
moon, flying loaded to a station in halo
orbit, and returning to the lunar
surface with tank empties plus a bit of cargo.
The orbit-to-orbit transport vehicle to LEO, is based at the station in the halo orbit
about the moon. It flies fully laden
from there to LEO, and returns to the
halo station with empties plus that same little bit of cargo. </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEghuINijB9GKf9_Ohjevms5UMe3vdL7alqCc18UGFQpwX-tW4Yxb_mIaSA-bxJeg9GCWKeK6nX2bq9hmd-Fs9t3vvxOg7MV8r2zycNu5GFFWbZcRN5E3DppminD0kh96piquGDUzBwGCKkaIFoOteRDvwW4WMGH8CI6-IQMXuSrbMWs6OCvxs5TWF5qugWU/s1031/f1%20basic%20notion.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="626" data-original-width="1031" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEghuINijB9GKf9_Ohjevms5UMe3vdL7alqCc18UGFQpwX-tW4Yxb_mIaSA-bxJeg9GCWKeK6nX2bq9hmd-Fs9t3vvxOg7MV8r2zycNu5GFFWbZcRN5E3DppminD0kh96piquGDUzBwGCKkaIFoOteRDvwW4WMGH8CI6-IQMXuSrbMWs6OCvxs5TWF5qugWU/w400-h243/f1%20basic%20notion.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Basic Notion of Lunar Propellant Manufacture for
Use in LEO<o:p></o:p></p>
<p class="MsoNormal">There are actually two halo orbit cases to consider, although overall, they are more-or-less a wash. One is the proposed halo orbit for NASA’s
“Gateway” station about the moon. This
one has an apoapsis radius that lies beyond the Hill sphere for orbit stability
about the moon. It also has a periapsis
altitude much higher than that used during Apollo, which acts to increase the delta-vee (dV)
required of any lander operating from that halo station. Long-term,
anything in this halo will eventually leave the moon and go into orbit
about the Earth. That is the direct result
of having an apoapsis distance outside the Hill sphere. <u>Thus the halo station requires periodic
correction burns</u>, <u>just to stay in
this orbit</u>.<o:p></o:p></p>
<p class="MsoNormal">Using LOX-LCH4 propulsion that does not push the
state-of-the-art as hard as SpaceX does with its Raptor engines, it takes almost 48 metric tons of propellant
manufacture on the lunar surface to deliver 1 metric ton of propellant to LEO
via the NASA halo station, and have the two
vehicles return to their bases. This
does <u>not</u> include the propellant necessary for orbital correction burns
to stabilize the station in this NASA halo orbit! The transport vehicle is the smaller for
this case, while the lander is the larger, as indicated in Figure 2. This unstable halo orbit choice may actually
be driven by the dV capability of SLS/Orion block 1 configuration, which cannot reprise the Apollo 8 mission. <o:p></o:p></p>
<p class="MsoNormal">The “recommended halo” eliminates entirely the need to ship
propellant to the halo station for periodic correction burns, precisely because it is stable. It has an apoapsis right at the Hill sphere
limit, and a periapsis at a distance
comparable to the old Apollo missions,
to reduce lander dV requirements.
It takes about 51 metric tons of propellant manufacture on the moon to
deliver 1 metric ton of propellant to LEO,
as also indicated in Figure 2.
For this case the transport is larger,
and the lander is smaller. <o:p></o:p></p>
<p class="MsoNormal">Neither of these halo-based options is more “economical”
than the projections for shipping propellant up to LEO from Earth’s surface, for SpaceX’s Starship/Superheavy vehicle, as indicated in the Figure. That vehicle also uses LOX-LCH4 propellants, so that comparison truly is “fair” in that
sense. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgcfTdKS5v2qPwj4u3TEg68zjMREFJjRrIQAuYLi4asctaX_PcXcBbGNvRAOzNccy9CZZSIKNGnzl0LHFjo_ImatqPKgOUJVKw9kjghALneLJwOyluM3Cmq8ieOgQ04mnEEO3rXFJCcFISb8SeAkiVCfXt1r-vuCEdeJcGngRmwXwcX78Y1QuicsBUBTY3R/s1031/f2%20overall%20result.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="626" data-original-width="1031" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgcfTdKS5v2qPwj4u3TEg68zjMREFJjRrIQAuYLi4asctaX_PcXcBbGNvRAOzNccy9CZZSIKNGnzl0LHFjo_ImatqPKgOUJVKw9kjghALneLJwOyluM3Cmq8ieOgQ04mnEEO3rXFJCcFISb8SeAkiVCfXt1r-vuCEdeJcGngRmwXwcX78Y1QuicsBUBTY3R/w400-h243/f2%20overall%20result.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Results for the Halo Orbit Cases<o:p></o:p></p>
<p class="MsoNormal">The alternative would be to manufacture propellant on the
lunar surface, and send it directly from
there to LEO, <u>without utilizing any
sort of halo orbit station as a waypoint</u>.
This does require entry from the lunar transfer orbit directly into a
low lunar circular orbit that is polar,
so that the south pole can be reached directly. The reverse is the return. Landings and takeoffs would be from-and-to
this low polar lunar orbit. <o:p></o:p></p>
<p class="MsoNormal">The dV requirements for the direct trip are higher than even
the halo transport dV’s. It was not
possible to get “reasonable” results with LOX-LCH4 propulsion for this
mission, the Isp levels are simply too
low. I had to resort to the higher Isp
of LOX-LH2 propulsion to get something “reasonable” in size. However,
since there may well be recoverable ice,
but no free carbon, on the
moon, it is far more likely that it will
be LOX-LH2 propellants that actually get manufactured there, anyway.
The corresponding results are given in Figure 3.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjXTPxcW-4kU9GDePGiWzEMmX2DZDnlpRU-PScA0AwQj9EhD5EWDZNYNKx7OdRJTzRSXgoIX-GJgI5hsLXOblUhBPFVqyUQQ-xMyU0x_WIS5Bz6xKpRK8HpOBOrnR1lMIp1UYAjj2k8csAKGTXLdSVC5xIzdg1K9qfuqg2NnUT6BKBnOVf3mNNql_z74-BM/s1034/f3%20final%20overall%20result.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="627" data-original-width="1034" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjXTPxcW-4kU9GDePGiWzEMmX2DZDnlpRU-PScA0AwQj9EhD5EWDZNYNKx7OdRJTzRSXgoIX-GJgI5hsLXOblUhBPFVqyUQQ-xMyU0x_WIS5Bz6xKpRK8HpOBOrnR1lMIp1UYAjj2k8csAKGTXLdSVC5xIzdg1K9qfuqg2NnUT6BKBnOVf3mNNql_z74-BM/w400-h243/f3%20final%20overall%20result.png" width="400" /></a></div><div class="separator" style="clear: both; text-align: center;"><u style="text-align: left;">Figure 3 – Results For Direct Delivery of Lunar
Propellant to LEO, Using LOX-LH2
Propulsion</u></div>
<p class="MsoNormal"><u>This is the only outcome that is better than trying to
ship propellant up from Earth’s surface to LEO</u>, even using a vehicle as capable as SpaceX’s
Starship/Superheavy. It takes 14 metric
tons of lunar propellant manufacture to support the delivery of 1 metric ton of
propellant from the moon to LEO. This
includes the vehicle returning all the way to a landing near the moon’s south
pole, with empties and some cargo. <o:p></o:p></p>
<p class="MsoNormal"><b>Outcomes<o:p></o:p></b></p>
<p class="MsoNormal"><span style="background: yellow; mso-highlight: yellow;">In
summary, <u>direct shipment from the
moon to LEO is the best option</u>, but
it will require the higher Isp of LOX-LH2 propulsion!</span> Using LOX-LCH4 is not feasible in any
“reasonable” vehicle sizes, primarily limited
by achievable mass ratio as you increase propellant. This requires manufacture on the moon of <span style="color: red;">25.1</span> metric tons of LOX-LH2 to deliver 1 ton of LOX-LH2 as payload to LEO. <span style="color: red;">(corrected)</span><o:p></o:p></p>
<p class="MsoNormal">Failing that, <u>shipping
propellant up from Earth’s surface to LEO is the next best option</u>, using the SpaceX Starship/Superheavy
vehicle, which is powered by LOX-LCH4
propulsion. This is projected to require
the manufacture of some 32 to 47 metric tons of LOX-LCH4 propellant on Earth to
deliver 1 metric ton of it (or instead a ton of LOX-LH2) to LEO. <o:p></o:p><span style="color: red; font-family: Calibri, sans-serif; font-size: 11pt;">This depends upon deliverable payload being 100-150 tons.</span></p>
<p class="MsoNormal"><u>The halo station options turned out to be the least
attractive</u>. Doing it with two
vehicles instead of one reduces payload delivered for the propellant used, by two vehicle payload fractions compounded
(because there are two vehicles),
instead of just one. The two halo
options are not much different at about 48 (NASA) and about 51 (recommended)
metric tons propellant to be manufactured on the moon to deliver 1 ton of
propellant to LEO. <u>The NASA halo
requires still more propellant manufacture on the moon than that about-48
figure</u>, because its halo station
requires correction burns just to stay in its orbit long term! The “recommended halo” avoids that
requirement, and gets a smaller lander
design, at the expense of a larger
transport vehicle to LEO. <o:p></o:p></p>
<p class="MsoNormal"><b>Conclusion<o:p></o:p></b></p>
<p class="MsoNormal">The final result says go with the higher-Isp LOX-LH2
propulsion, and operate direct from the
lunar surface to LEO fully laden, and back to the lunar surface very lightly
laden. Return trips return the empty
tanks for the next propellant shipment,
plus in this study, a couple of
tons of payload that could be base operating supplies.<o:p></o:p></p>
<p class="MsoNormal"><b>Be Aware<o:p></o:p></b></p>
<p class="MsoNormal">The dV requirements used in this study include 8% of
midpoint speed as dV budgets for course correction, and 0.2 km/s budgets for rendezvous and
docking, once close (within at most 10
km). Lunar takeoff ideal dV values were
factored-up by 1.0083 for gravity losses,
and lunar landing ideal dV values were factored-up by 1.50 for losses
plus the dominant hover and divert budget requirements. Everything else was presumed loss-free “impulsive”, for factor 1.00 applied to ideal values from
basic 2-body orbital mechanics. The
3-body approach and departure problems were approximated by the “far V” versus
“near V” energy approximation. <o:p></o:p></p>
<p class="MsoNormal"><b>Explanations<o:p></o:p></b></p>
<p class="MsoNormal">There are some fundamental trends of mass ratio capability
and the dV it can produce, vs added
propellant and Isp, which help explain
these results. These are depicted
generically in Figure 4. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEipyy_qILo5g-PDyfys_uRasOhWOd2lA8bI8BBVPRFBzlE4OCIcTjX4Wnp8WAVpJC-DeqGReF87Vz111tOnwUG5DPgimnpWTLW8Z1vxuyfzxRiSVju2zEi2M7tZSNBtl2IeJIoNo736qoaAWmR8CZDqJ4a21pVqr2N67vAEOl4NkL1NED0hUNR2saweRqZb/s1034/f4%20trends.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="627" data-original-width="1034" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEipyy_qILo5g-PDyfys_uRasOhWOd2lA8bI8BBVPRFBzlE4OCIcTjX4Wnp8WAVpJC-DeqGReF87Vz111tOnwUG5DPgimnpWTLW8Z1vxuyfzxRiSVju2zEi2M7tZSNBtl2IeJIoNo736qoaAWmR8CZDqJ4a21pVqr2N67vAEOl4NkL1NED0hUNR2saweRqZb/w400-h243/f4%20trends.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Trends That Explain Results <o:p></o:p></p>
<p class="MsoNormal">The “knees” in these curves are not so apparent in the MR vs
added Wp plot, but they are very
apparent in the dV vs added Wp plots parametric on Isp. Where the slope is steep, you will get a better (lower) propellant
burned vs payload delivered ratio. Where
the slope is shallow, that ratio will be
large and unfavorable. <o:p></o:p></p>
<p class="MsoNormal">Basically, there is a
favorable range of deliverable dV at each Isp level. For the 450 s Isp level typical of LOX-LH2
propulsion, this is up to 8 km/s dV for
really nice results, and up to about 11
km/s at the very most. Beyond that, it is a very serious diminishing-returns
problem: adding a lot of propellant for
almost no improvement. <o:p></o:p></p>
<p class="MsoNormal">For the 350-400 s Isp level typical of LOX-LCH4
propulsion, the most favorable range is
up to about 7 km/s dV. At the most, it is about 9 km/s. Beyond that,
this is pretty much pointless. <o:p></o:p></p>
<p class="MsoNormal">By switching the halo station based lander and transport
designs to LOX-LH2 (instead of LOX-LCH4),
the propellant to payload ratios could be significantly reduced, perhaps looking more like those of Earth
launch with Starship/Superheavy, or
maybe even slightly better. <o:p></o:p></p>
<p class="MsoNormal">Regardless, the lunar
surface based and launched single direct transfer design approach is still the
best, despite it being only marginally
favorable on the dV vs Wp curve. That is because it is a single vehicle, and not two vehicles, as in the other scenarios. This scenario would look even better if its
propulsion were nuclear thermal at 700+ s of Isp. <o:p></o:p></p>
<p class="MsoNormal">As it is with LOX-LH2 propulsion, the total vehicle dV requirement could be
reduced a little, making the propellant
used/propellant delivered ratio even better,
if an LEO tug were used to retrieve the vehicle from an elliptical
capture orbit about the Earth. The same
tug could put the vehicle back into the elliptical orbit for departure. That reduces the arrival and departure dV’s
significantly, and it eliminates the
rendezvous and docking dV requirement.
This gain is largely offset by the need for propellant deliveries to
power the tug, though.<o:p></o:p></p>
<p class="MsoNormal"><b>Caveat<o:p></o:p></b></p>
<p class="MsoNormal">Ullage solutions for multiple burns with cryogenic liquid
propulsion were NOT determined for any of these design rough-outs. But they will have to be, to flesh out all the design requirements! Attitude control was also not addressed, although given adequate acceleration
levels, some of those thrusters could
supply the ullage function. That is
determined by the settling time constants that are acceptable. <o:p></o:p></p><p></p><p class="MsoNormal"><b><span style="color: red;">Corrections 12-10-2023:<o:p></o:p></span></b></p>
<p class="MsoNormal"><span style="color: red;">I had not followed through fully on
the spreadsheet for the direct vehicle.<span style="mso-spacerun: yes;">
</span>The 14:1 delivery ratio figure goes with an otherwise-converged design
that had far-insufficient thrust to takeoff and land,<span style="mso-spacerun: yes;"> </span>resulting in too low an inert mass.<span style="mso-spacerun: yes;"> </span>When I corrected that,<span style="mso-spacerun: yes;"> </span>the vehicle proved to be enormous at 3000
tons,<span style="mso-spacerun: yes;"> </span>with a really bad-looking delivery
ratio.<span style="mso-spacerun: yes;"> </span>I reconverged multiple times with
multiple candidate engine numbers and thrusts,<span style="mso-spacerun: yes;">
</span>until all the gees looked good,<span style="mso-spacerun: yes;"> </span>including
landing with hover capability,<span style="mso-spacerun: yes;"> </span>but with
takeoff reduced to 0.5 gees over lunar surface gravity.<span style="mso-spacerun: yes;"> </span>That got me to the corrected figure. <o:p></o:p></span></p><br /><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com2tag:blogger.com,1999:blog-2675974463524895416.post-16615668254035015022023-11-22T09:58:00.000-06:002023-11-22T09:58:31.572-06:00How the Suborbital “Hopper” Calculations Were Made and with What<p>The Mars rocket hopper design rough-out was done using the
course materials and tools for the “Orbit Basics +” course offered on the New
Mars forums. There is an “orbit basics”
spreadsheet that does elliptical orbit 2-body calculations for either the <u>two-endpoints
case</u> or the <u>R-V-q observation case</u>.
That tool’s R-V-q option can create suborbital trajectories, which was done for the rocket hopper. The spreadsheet calculates speeds V at
periapsis, apoapsis, and at any one user-input radius. <b>See Fig.
1.</b> </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj4Yz2PMxUgi1u9NjRZMVOI6_k3FTBH5Rs6QX0dbIskk0amGWaaotwrYh1woHPrBCvDFSIKmaUZyHRwtFZU2qZ72RwWswYt0nfUpeFCRRbfxJJH-arPhnBhQsGW042d9j_GCg1MKBeVIYLmcm161fSFqs53C6gn2BreJEmWsL1OS3VqUnKEFqZtkFtJ0U8e/s998/F1%20spreadsheet%20orbits%201.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj4Yz2PMxUgi1u9NjRZMVOI6_k3FTBH5Rs6QX0dbIskk0amGWaaotwrYh1woHPrBCvDFSIKmaUZyHRwtFZU2qZ72RwWswYt0nfUpeFCRRbfxJJH-arPhnBhQsGW042d9j_GCg1MKBeVIYLmcm161fSFqs53C6gn2BreJEmWsL1OS3VqUnKEFqZtkFtJ0U8e/w400-h240/F1%20spreadsheet%20orbits%201.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – The Two Cases Handled by the Orbit Basics
Spreadsheet<o:p></o:p></p>
<p class="MsoNormal">To use this tool for the rocket hopper, the most effective way was to define an exit
(and by symmetry entry) point at the edge of Mars’s atmosphere, and investigate various speeds V and exit
angles relative to local horizontal. Not
every combination is allowable, only
certain values produce survivable peak heating and peak deceleration gees, and also a feasible end-of-hypersonics
altitude, for a direct rocket-braked
landing. In fact, many combinations produced instead a surface
impact while still quite hypersonic, in
Mars’s thin atmosphere. <o:p></o:p></p>
<p class="MsoNormal">The symmetry of the exposed portion of the ellipse makes the
V and angle “a” values the same for exit and entry, at the entry interface altitude. That is exactly how the suborbital trajectory
analysis links directly to the hypersonic entry analysis. <o:p></o:p></p>
<p class="MsoNormal">For the launch speed required of the hopper, we need the speed along the orbit at the
surface of the planet. We need to be
moving that fast at just about the same angle “a”, at the end of the launch burn. That is the theoretical dV<sup>o</sup> value, which needs to be factored up by about 1.02
to cover gravity and drag losses on Mars.
The factored-up launch dV is the mass ratio-effective value needed for
proper use in the rocket equation. <b>See
Fig. 2.</b><o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi78-pU1uqJqKVVJcdHtHDTeAbuWjJpDJjmATaO1zkf48DoJGE_l7zwvpf6sN4uVS3Qnnw3Scl4qGlZk7GHxnyVy9oxOqTtvBZRmZwna3YAPscQknK3hHWVWEYZUGEp1yP0T4HRG2YZN5ePuix68gl_5l00F33n95YTXlJHbH444CDFCOw9xfalSrZLjLFW/s998/F2%20spreadsheet%20orbits%202.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi78-pU1uqJqKVVJcdHtHDTeAbuWjJpDJjmATaO1zkf48DoJGE_l7zwvpf6sN4uVS3Qnnw3Scl4qGlZk7GHxnyVy9oxOqTtvBZRmZwna3YAPscQknK3hHWVWEYZUGEp1yP0T4HRG2YZN5ePuix68gl_5l00F33n95YTXlJHbH444CDFCOw9xfalSrZLjLFW/w400-h240/F2%20spreadsheet%20orbits%202.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Using the R-V-q Option in the Orbit Basics
Spreadsheet for Suborbital Trajectories<o:p></o:p></p>
<p class="MsoNormal">The ”Orbits +” course covers launch, entry,
descent-and-landing, use of the
rocket equation, and estimating real
engine performance, as well as 2-body
orbital mechanics of elliptical orbits. For the rocket hopper, both <u>entry</u> and <u>descent-and-landing
apply</u>, using the methods and tools
that are part of the course. The direct
rocket-braked landing is so simple, it
can be estimated from hand calculations. <o:p></o:p></p>
<p class="MsoNormal">The entry analysis is a 2-D Cartesian simplified analysis
dating to about 1953, and attributed to
H. Julian Allen. It was used in the
1950’s for estimating entry conditions for ICBM and IRBM warheads. It was declassified by the mid-1960’s, and then taught in engineering school
classes. Entry is presumed to happen
along a straight line trajectory at a fixed entry angle. The range is a crude estimate that you must
wrap around the curved surface of the planet.
The constant angle you have to presume is relative to local
horizontal, as you move around the curve
of the planet’s surface. <o:p></o:p></p>
<p class="MsoNormal">These crude estimates get you “into the ballpark” only! There is no substitute for a real digital
trajectory program in polar coordinates,
but you do have to expend the significant efforts to construct the model
to run in it. <b><i>At this stage of the
game, that is very inconvenient, since the model to be input changes
drastically as you iterate configurations.
Hence the need for a quicker ballpark estimate.<o:p></o:p></i></b></p>
<p class="MsoNormal">There is a lesson in the “Orbits +” course that deals with
using the simplified entry analysis as a spreadsheet model of the entry process. That spreadsheet is supplied as part of the
course materials. <o:p></o:p></p>
<p class="MsoNormal">In reality, there is
significant trajectory “droop” after the peak deceleration gees point, that the simplified analysis does not account
for. I merely presume the local angle
has increased to 45 degrees down, by the
Mach 3 end-of-hypersonics point, when I
do the rocket-braking by-hand calculations.
<o:p></o:p></p>
<p class="MsoNormal">There is also a lesson in the “Orbits +” course that deals
with multiple ways to land after the hypersonics are over. There is no spreadsheet, but all the calculation equations are there
to estimate any of these things by hand.
For the thin atmosphere of Mars,
from inevitable very low end-of-hypersonics altitudes with multi-ton
vehicles, there really is only direct
rocket braking as a feasible thing to do.
<o:p></o:p></p>
<p class="MsoNormal">There is no time to deploy a chute, much less get any deceleration from it, plus there are no chute designs capable of
surviving opening at Mach 3. Even the
ringsail chute designs used for probes at Mars have a maximum opening speed of
Mach 2.5, and slower-still is preferred
as more reliable. <o:p></o:p></p>
<p class="MsoNormal">Direct rocket braking is actually the simplest case, and easily figured with nothing more than the
simple kinematics of a high school-level physics course. <b>See Fig. 3.</b> <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjEVaoxS_ZsTvvyXiYqp94gbyG2hjzQ0ZhD-GSSfCm5EtHn_kkpZC7X2nz5zSobA2-WUj_LER8TaIKlPGqZGtgIKOjcySj6nJwg3Zjj9XTcalWEzYyomxpjq8WVHVphgmGGkw2lsWYXwSOvG0mTsZa6yU663iUK65ltosDPvBLHyEVs-sAcvexcIQ8h5N8O/s998/F3%20entry%20descent%20and%20landing.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjEVaoxS_ZsTvvyXiYqp94gbyG2hjzQ0ZhD-GSSfCm5EtHn_kkpZC7X2nz5zSobA2-WUj_LER8TaIKlPGqZGtgIKOjcySj6nJwg3Zjj9XTcalWEzYyomxpjq8WVHVphgmGGkw2lsWYXwSOvG0mTsZa6yU663iUK65ltosDPvBLHyEVs-sAcvexcIQ8h5N8O/w400-h240/F3%20entry%20descent%20and%20landing.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – The Entry Model,
Plus Descent-and-Landing for Direct Rocket Braking<o:p></o:p></p>
<p class="MsoNormal">The vehicle layout and dimensions, plus its weight statement, are essentially custom hand
calculations, the suite of which is
different for each different configuration class. I started with three configurations, but only one gave me the low ballistic
coefficient that the entry analyses said I must have. I included wide-stance folding landing legs
for rough-field operations.
Clearly, there are a lot of
considerations to address. I created a
custom spreadsheet to estimate all these quantities rapidly, since I had to iterate multiple times before
identifying a feasible solution.<o:p></o:p></p>
<p class="MsoNormal">The “Orbits +” course has a lesson on vehicle layout, and a spreadsheet by which to set the weight
statement, but that spreadsheet was not
really suitable for this very specialized suborbital vehicle, especially since it must enter the
atmosphere, and also do that entry dead-broadside
to get the necessary lower ballistic coefficient. It is critical to select the correct diameter
for this kind of vehicle, so that the
lengths are in the correct range, and
those results must be compatible and consistent with the seating arrangements
in the passenger cabin. That’s why I did
it as a custom calculation, and why I
created my own spreadsheet for that purpose.
<b>See Fig. 4.</b> <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1DIfxjUBPT5aup4V_OZfaIkUZQCXgq2VZu2FdfnlJPsI4tCCIwCI1vA5CN5nGoXLqblXOVcnIwciT5lNfyp1kpe2MaymT-eoChtAL1Q35YNZCHI9-HC4RUlwED7n3OiJ4KpBat_lmlVTl7C2PWQ0wLwq3EAFCouqTiKYobvQyLsOFNeNTEAxO7Ae1R5Bu/s998/F4%20veh%20layout%20considerations.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1DIfxjUBPT5aup4V_OZfaIkUZQCXgq2VZu2FdfnlJPsI4tCCIwCI1vA5CN5nGoXLqblXOVcnIwciT5lNfyp1kpe2MaymT-eoChtAL1Q35YNZCHI9-HC4RUlwED7n3OiJ4KpBat_lmlVTl7C2PWQ0wLwq3EAFCouqTiKYobvQyLsOFNeNTEAxO7Ae1R5Bu/w400-h240/F4%20veh%20layout%20considerations.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Downselecting to One Configuration for Vehicle
Layout<o:p></o:p></p>
<p class="MsoNormal">All of this is aimed at using the rocket equation to relate
vehicle weight statement to its velocity-increment (dV) performance
capability. The spreadsheet in the
lesson on vehicle sizing of the “Orbits +” course does exactly that, in a spreadsheet that is supplied as part of
the course materials. Since I did the
hopper with a custom layout sheet, I had
to include this rocket equation stuff in it.
<o:p></o:p></p>
<p class="MsoNormal">The classic rocket equation dV = Vex LN(MR) uses the vehicle
weight statement (from a vehicle layout process) to determine mass ratio MR =
Wign/Wbo, and an estimate of engine Isp
to determine the effective exhaust velocity Vex = Isp * gc. It then gives you the performance estimate
dV, which must cover the mission needs
plus any gravity and drag losses, or
other considerations, such as hover and
divert during landings. <o:p></o:p></p>
<p class="MsoNormal"><u>There is a restriction on this</u>: you may sum the dV values estimated for all
the mission burns into an overall mission dV,
<u>only</u> if the weight statement <u>does not change between burns</u>. That means the payload and inert masses do
not change, and the only propellant mass
changes are those for the burns. Failing that restriction, you have a slightly different weight
statement each time one of those items changes.
You must do a separate rocket equation calculation for only the burn
associated with each slightly-different weight statement. This hopper does not change its weight
statement between burns!<o:p></o:p></p>
<p class="MsoNormal">For sizing vehicles,
the reverse process is what we really want to do, for which the rocket equation rearranges to
MR = exp(dV/Vex). The engine Isp
estimate gets us a Vex as before. The
mission dV is as before. <u>The layout
gets us a payload mass and an estimate of vehicle inert mass fraction</u>. We use the rocket equation in reverse with
the mission dV and the engine Vex to determine the MR that is required. <o:p></o:p></p>
<p class="MsoNormal">This MR result determines the propellant mass fraction = 1 –
1/MR. The payload fraction is 1 –
propellant fraction – inert fraction.
Payload divided by payload fraction is the ignition mass, ignition mass times the inert fraction is the
inert mass, and propellant fraction
times ignition mass is the propellant mass.
Payload plus inert is burnout mass,
and burnout plus propellant is ignition mass. <u>In effect,
we are finding the vehicle weight statement from mission dV and engine
performance to complete the vehicle layout process</u>. <b>See Fig. 5.</b><o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjMdHENSe8iNO8LmSDOElgMrENHupYrYuLJxFr-MhD3PBloXr6Wn8iOZEfelAxrNiFxg1QRubJt4TfXfYxlQz0wW4qWSRY28sORxaBp5ju1WiKxlEZBglbo1UT11UUY7W3N0YU_XdOneJqOfDc06nEyW7gXVkIyTG7iFCdP8-Scy_KrQ2wjdy8dvyPSDlYB/s998/F5%20sizing%20the%20vehicle.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjMdHENSe8iNO8LmSDOElgMrENHupYrYuLJxFr-MhD3PBloXr6Wn8iOZEfelAxrNiFxg1QRubJt4TfXfYxlQz0wW4qWSRY28sORxaBp5ju1WiKxlEZBglbo1UT11UUY7W3N0YU_XdOneJqOfDc06nEyW7gXVkIyTG7iFCdP8-Scy_KrQ2wjdy8dvyPSDlYB/w400-h240/F5%20sizing%20the%20vehicle.png" width="400" /></a></div><p class="MsoNormal">Figure 5 – Using the Rocket Equation Properly to Size
Vehicles to Missions<o:p></o:p></p>
<p class="MsoNormal">Clearly, an accurate
estimate of expected engine performance (as Isp or Vex) is crucial to the
results! There are a lot of references
out there that list tables of Isp versus propellant combinations. Just picking one right out of such tables is
a serious error! That is because engine
Isp depends at least as much on the nozzle expansion characteristics, as it does the propellant combination. The expansion in the table is rarely the one
you want to use, and nozzle efficiency
effects are <u>never included</u> in those tables. <o:p></o:p></p>
<p class="MsoNormal">These things are all functions of the chamber pressure, as measured at the nozzle entrance. <u>The chamber pressure value used in the
tables is rarely the value you want to use</u>.
<o:p></o:p></p>
<p class="MsoNormal">Finally, Isp is
directly affected by the engine cycle (through the dumped bleed gas fraction), which those tables <u>never include</u>. You can easily be 10%-or-more wrong just
pulling values out of those tables. Due
to the exponential nature of the rocket equation, that error in Isp can lead to fatal errors in
your vehicle results for mass ratio and weight statement. <o:p></o:p></p>
<p class="MsoNormal">Thrust is often represented in terms of chamber pressure as
F<sub>th</sub> = C<sub>F</sub> Pc At.
Isp is thrust divided by flow rate,
<u>but it has to be the flow rate drawn from the tanks to be consistent
with the rocket equation</u>. Flow rate
from tanks = flow rate through nozzle + flow dumped overboard. The flow rate through the nozzle relates to
chamber pressure and c*-velocity as Pc C<sub>D</sub> At g<sub>c</sub> /
c*. And c* is a weak power function of
Pc, where the exponent is usually in the
vicinity of 0.01. The specific heat
ratio of most rocket gases is in the vicinity of 1.20. <b>See Fig. 6,</b> for which the <u>only</u> propellant
combination-related item is c*.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgw4H2itICp3_-fvhhffADeLV4wANwe0emdnMvXGmcKBxshZMTlTn9kyvkzue6vAAlFPWLglb-NCPPsGdH2SsEZ29jtKnfR24VoutEc3RZjA9TQBnj6uzAtEinfiJF_h-0o2nnDnVVRFOvzl_SXB1wIrqPjb7H3yijDJMVIwA0-dRTGwxu3sJNQFdL2tLqe/s998/F6%20estimating%20Isp.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="600" data-original-width="998" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgw4H2itICp3_-fvhhffADeLV4wANwe0emdnMvXGmcKBxshZMTlTn9kyvkzue6vAAlFPWLglb-NCPPsGdH2SsEZ29jtKnfR24VoutEc3RZjA9TQBnj6uzAtEinfiJF_h-0o2nnDnVVRFOvzl_SXB1wIrqPjb7H3yijDJMVIwA0-dRTGwxu3sJNQFdL2tLqe/w400-h240/F6%20estimating%20Isp.png" width="400" /></a></div><p class="MsoNormal">Figure 6 – How Engine Performance Must Really Be Estimated
for a Specific Design<o:p></o:p></p>
<p class="MsoNormal">You are not totally free to set an arbitrary expansion ratio
Ae/At! It does not matter whether your
nozzle is a “sea level” design or a ”vacuum-adapted” design, any engine that is to be tested in the open
air at sea level on Earth must not be allowed to flow-separate, because that risks destruction of at least
the nozzle exit bell! <u>Testing into a
vacuum tank is extremely expensive</u>!<o:p></o:p></p>
<p class="MsoNormal">For any given expanded pressure in the exit plane, there is a value of the ambient atmospheric
“back pressure” Pback that is “too much”,
causing flow separation. That
level is denoted Psep, and it is easily
estimated from the nozzle expansion pressure ratio: Psep/Pc = (1.5 Pe/Pc)<sup>0.8333</sup>, which is an entirely empirical correlation
developed for conical nozzles, and is
only slightly conservative for curved bells.
<o:p></o:p></p>
<p class="MsoNormal">For a “sea level” nozzle design, you want predicted Psep = sea level
barometric, at some part-throttle
Pc. That way, you can test in the open air for all power
settings that high, or higher. The same is true of “vacuum-adapted”
designs, unless you give up testing in
the open air! But even then, the engine and its nozzle still have to fit
within the allotted space behind the stage. <o:p></o:p></p>
<p class="MsoNormal">The “Orbits +” course has a lesson on this topic, plus a spreadsheet tool that does all these things. It includes a database of c* and r-value data
versus several propellant combinations,
as functions of Pc.<o:p></o:p></p>
<p class="MsoNormal"><b>Updated 11-21-2023:</b>
These very same methods were used to compute revised data for the upgraded
Mars rocket “hopper” that could also serve as a personnel taxi to low Mars
orbit. <o:p></o:p></p>
<p class="MsoNormal">The original suborbital rocket “hopper” design summary was
posted on this site as “Rocket Hopper for Mars Planetary Transportation”, dated 1 November 2023. The upgraded “hopper” that can also serve as
an orbital taxi is posted on this site as “Upgraded Rocket Hopper as Orbital
Taxi”, dated 21 November 2023. <o:p></o:p></p>
<p class="MsoNormal">There is a completely unrelated posting that deals with
long-distance bulk freight transport on the surface of Mars. That one is “Surface Freight Transport on
Mars”, dated 4 November 2023. <o:p></o:p></p>
<p class="MsoNormal">The final landing choice not described here is the lifting
pull-up proposed by SpaceX for landing its Starship vehicle on Mars. That is distinct from direct rocket
braking, and from parachute-assisted
descents, which require terminal rocket
braking on Mars. It is covered in the
entry, descent, and landing lesson of the “orbits +” course
materials. <o:p></o:p></p>
<p class="MsoNormal">I did not examine that choice for any of these rocket
“hopper” designs, because I did not
believe that my cylindrical layout has the mild-supersonic lift/drag ratio
necessary to execute an aerodynamic pull-up,
even at very low altitudes on Mars.
I don’t really believe SpaceX’s Starship can do that either, but that would be another study. <o:p></o:p></p>
<p class="MsoNormal">To access the “orbits +” course materials, which includes the spreadsheets, go to the Mars Society’s New Mars forums
online. Go to the “Acheron Labs”
section, “interplanetary transportation”
topic. On about the second page of the
list of conversation threads, <b><i>look
for the “orbital mechanics class traditional” thread.</i></b> The course materials are actually posted elsewhere
online, <u>but the links to each class
session’s materials are in posts 3-to-21 of that thread</u>. <o:p></o:p></p>
<p class="MsoNormal">You will have to download the Excel spreadsheet files to
make them functional. The classes have a
sort of lecture session (numbered) and a problem-working session (numbered with
a “B” suffix). These are available as
Powerpoint slide sets and as pdf documents that are basically the
traditional-style textbooks. I recommend
you download the pdf textbooks, because
all the explanations are in there. They
would be partly missing in the slide sets. <o:p></o:p></p><p><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-53968956711715735132023-11-21T11:52:00.002-06:002023-11-22T12:32:37.376-06:00Upgraded Rocket Hopper as Orbit Taxi<p class="MsoNormal">This article is about modifying a pre-existing design
rough-out for a suborbital Mars rocket “hopper”, into a design also capable of operating as a
low Mars orbit personnel taxi. That
original rocket hopper design rough-out is covered in the article titled
“Rocket Hopper for Mars Planetary Transportation”, dated 1 November 2023, and posted on this site. <o:p></o:p></p>
<p class="MsoNormal"><o:p> </o:p></p>
<p class="MsoNormal"> <b><u>The
Problem<o:p></o:p></u></b></p>
<p class="MsoNormal">Started with a suborbital “hopper”<o:p></o:p></p>
<p class="MsoNormal"> 10
persons aboard on p-suits<o:p></o:p></p>
<p class="MsoNormal"> Short-term
life support plus small luggage<o:p></o:p></p>
<p class="MsoNormal">Could it also serve as a low orbit taxi?<o:p></o:p></p>
<p class="MsoNormal"> Same
payload<o:p></o:p></p>
<p class="MsoNormal"><b> </b></p>
<p class="MsoNormal">As indicated in the table just above, I started with the earlier design rough-out
that was only a suborbital “hopper”. The
idea was to carry 10 persons as the payload.
Although the cabin is pressurized,
these persons ride in pressure suits for a safety backup. There are limited supplies of oxygen and
drinking water, plus minimal snack
foods, for up to a few hours’ ride. A small luggage allowance was included. The same payload would be carried to any low
orbit destination.<o:p></o:p></p>
<p class="MsoNormal">As indicated in <b>Figure 1</b> just below, the suborbital trajectory is actually an
ellipse in polar coordinates, one with
its periapsis inside the planet. The
vehicle launches into a gravity turn that reaches a suitable velocity and path
angle at the entry interface altitude,
coasting from there. <o:p></o:p></p>
<p class="MsoNormal">The best place to do a course correction is the apoapsis
outside the sensible atmosphere, where
speeds are lowest and directions are easiest to change. The entry conditions mirror the exit
conditions, with no burn. The landing is a direct rocket-braked descent
from the end-of-hypersonics point at local Mach 3 (about 0.7 km/s speed). 45
degrees of trajectory “droop” along a straight-line path is presumed. I factored-up the speed to “kill” by 2, to budget the final landing mass
ratio-effective delta-vee (dV). <o:p></o:p></p>
<p class="MsoNormal">As illustrated in <b>Figure 2 </b>just below, I used a surface-grazing ellipse as the
initial transfer trajectory to the 300 km nominal low orbit altitude. Like the long-range suborbital mission, the vehicle launches into a gravity
turn, putting it onto the proper path at
the entry interface altitude, at end of
launch burn. Only the path angle is
different, being a lot smaller. The entry point after de-orbiting is the
mirror image. <o:p></o:p></p>
<p class="MsoNormal"><o:p></o:p></p><p>There is a small burn at apoapsis to raise the periapsis to
the entry interface altitude, with a
period shorter than the target low circular orbit altitude. This ellipse is the parking orbit in which to
“chase” any target in the low circular orbit.
Once synchronized, there is
another small burn to circularize,
followed by a traverse to rendezvous,
plus a budget to actually dock.
Deorbiting is another small burn,
back onto the surface-grazing ellipse that guarantees entry. The direct rocket-braked landing is identical
to that of the long suborbital trajectory,
except that, as it turned
out, the end-of-hypersonics altitude is
higher, coming back from orbit at the
lower entry angle. </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhmdjcCJRN7NbiT3Tq6DodOIzZ5JF9Peweo_iSNtjg6Yz-_UQcZAp3XP4gzfRaN1UWmqYVqL2zg19l9kLBgNwVtW_IFGh05ZyhBQuajSDSoyn91opUw6B3XAlfvXO6ES3ZwpyfzMBZT_zwvh-8vuA2AlrM3FfHsrzAX-ahvBGI-oD024eiRpUNiIQUPQfEY/s1017/f4%20suborbital%20dV.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="1017" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhmdjcCJRN7NbiT3Tq6DodOIzZ5JF9Peweo_iSNtjg6Yz-_UQcZAp3XP4gzfRaN1UWmqYVqL2zg19l9kLBgNwVtW_IFGh05ZyhBQuajSDSoyn91opUw6B3XAlfvXO6ES3ZwpyfzMBZT_zwvh-8vuA2AlrM3FfHsrzAX-ahvBGI-oD024eiRpUNiIQUPQfEY/w400-h240/f4%20suborbital%20dV.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Suborbital Missions, Longest-Range Shown <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgRO5b-CfRk6THe37zBspsl8LktptgJMx4aZu4V3Tr45xSfntnZ_CI8KYJ5T_hcX1Q5fT3BXco-7WwOcjR7d0nAoV8lPWl3KIGj-J4Z7gc6rBQjU-hUioEiFESTSw0W8N6WDJ8WDGRojcnwMzf5qCHiqjlW5ICpve5u9b3rHfNRyF9SX31QrC1iXhqPZBWo/s1012/f3%20mission%20dV.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="621" data-original-width="1012" height="245" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgRO5b-CfRk6THe37zBspsl8LktptgJMx4aZu4V3Tr45xSfntnZ_CI8KYJ5T_hcX1Q5fT3BXco-7WwOcjR7d0nAoV8lPWl3KIGj-J4Z7gc6rBQjU-hUioEiFESTSw0W8N6WDJ8WDGRojcnwMzf5qCHiqjlW5ICpve5u9b3rHfNRyF9SX31QrC1iXhqPZBWo/w400-h245/f3%20mission%20dV.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – The Orbital Mission, Including “Chase”, Rendezvous,
and Docking <o:p></o:p></p>
<p class="MsoNormal">To accommodate the more demanding mission, I resized the candidate LOX-LCH4 engine
design, and revised the inert masses
upward a little. Entry conditions forced
me to increase the diameter and length a little, in order to keep the entry ballistic
coefficient down to tolerable values.
The original rough-out had two sets of tanks: mains and headers. The landing and course correction propellant
was in the headers, with the launch
propellant in the mains. <o:p></o:p></p>
<p class="MsoNormal">This became 3 sets of tanks and two different engine designs. The launch-and-landing main engines stayed
about the same at 30,000 lb thrust, each
of 4, drawing from the mains for launch
and headers for landing. I was able to
increase the expansion ratio and specific impulse a little bit. See <b>Figure 3</b> for the basic layout
revisions.<o:p></o:p></p>
<p class="MsoNormal">But course correction suborbitally, and all the orbital maneuvering, rendezvous,
and docking, really needed much
lower thrust levels. So I sized some
lower-pressure, pressure-fed engines of
only 550 lb thrust, each of 4. These used a small third set of 800 psi
pressurized propellant tanks, plus a
supply of dry nitrogen gas at 2200 psi to power this, in one of two options examined. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgJlovOi5-5fTHLGa8X44HO8qHKk9feY8jHVpAUUPC1qX5WiUVgTQ_bCrzd178TRsCdzBtnUxJtnqZ5BxryUjbX2oIr5I9hqM1Tvd9gC-ya2DuykpqC6zjKiOucOkvA8svkVWIZta_jlwzJlUho6JMnvY99Jsc-KIakeEM0x2DQtbBnGCTzygftRr6Aucch/s977/f18%20revised%20layout.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="591" data-original-width="977" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgJlovOi5-5fTHLGa8X44HO8qHKk9feY8jHVpAUUPC1qX5WiUVgTQ_bCrzd178TRsCdzBtnUxJtnqZ5BxryUjbX2oIr5I9hqM1Tvd9gC-ya2DuykpqC6zjKiOucOkvA8svkVWIZta_jlwzJlUho6JMnvY99Jsc-KIakeEM0x2DQtbBnGCTzygftRr6Aucch/w400-h243/f18%20revised%20layout.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – Revised Internal Layout at Larger Diameter and
Length <o:p></o:p></p>
<p class="MsoNormal">Because the inert fraction increased a bit, I resized the expansion of the main engines
to increase specific impulse a bit, to
compensate as much as possible. The
original “hopper” main engines had an expansion ratio sized for incipient
separation at 67% chamber pressure, if
fired in the open air at sea level on Earth.
I raised that to 80%. See <b>Figure
4</b> just below. <o:p></o:p></p>
<p class="MsoNormal">The idea was to enable easy and relatively inexpensive
development testing on Earth. The change
was small, but every little bit
helps. These being full flow cycle, turbo-pumped engines of significant thrust
level, I did not want to complicate
things by adding vacuum bell extensions that were not regeneratively cooled. These do not push the state of the art very
hard, being only 2000 psia chamber
pressure.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjFQZnWY-gB-O1ylVwe8NC52N29Or9n1uwLZKwF2brjkIuccJaB_KBMFAS-GqshLIVxVckIXzHCQfk0RDrVykftp70aXqTem3QrroMa7QQF_hyphenhyphenaKQKBoFSdcH7wBDfE0KM6BInMk34HzRCklNlLkQDLGQCyIgpGXO0XyY7x_mG_u7D9LPrOZlpL5A7aqTRl/s1017/f10%20main%20engines.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="1017" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjFQZnWY-gB-O1ylVwe8NC52N29Or9n1uwLZKwF2brjkIuccJaB_KBMFAS-GqshLIVxVckIXzHCQfk0RDrVykftp70aXqTem3QrroMa7QQF_hyphenhyphenaKQKBoFSdcH7wBDfE0KM6BInMk34HzRCklNlLkQDLGQCyIgpGXO0XyY7x_mG_u7D9LPrOZlpL5A7aqTRl/w400-h240/f10%20main%20engines.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Reworked Main Engine Design for Slightly Higher
Expansion <o:p></o:p></p>
<p class="MsoNormal">The original “hopper” design study convinced me that I did
not need the large main engine thrust levels to do course corrections on the
suborbital missions, or orbital
maneuvering, rendezvous, and docking,
on the orbital mission. I kept
the redundancy of 4 engines, but sized
for crudely only 0.1 gee of vehicle acceleration, once exoatmospheric. <o:p></o:p></p>
<p class="MsoNormal">Since the propellant quantities would be small, the simplification of a pressure-fed design would
be beneficial. Alternatively, since the engines were small, they could be fed by electric-driven
positive-displacement pumps. Either
way, I picked a simple conical bell shape, two-piece,
with a bell extension that was not regeneratively cooled, as shown in <b>Figure 5 </b>just below. Development testing on Earth could be done
without the extension, but operations on
Mars or in space would use the bell extension.
This was not a throttleable design. <o:p></o:p></p>
<p class="MsoNormal">I ran numbers both ways for the propellant feed to the
maneuvering engines. I did not like the
pumping power required for the positive-displacement pumped option. It implied very heavy batteries, even for the modest propellant
quantities. Regulated constant inert gas
pressure on the propellant tanks turned out to be the better option. These used a small third set of 800 psi
pressurized propellant tanks, plus a
supply of dry nitrogen gas at 2200 psi to power this. The chamber pressures were low enough to keep
the pressures fairly modest on the tanks,
so that at small size, they were
not that big an inert weight penalty. See
<b>Figure 6</b> just below. <o:p></o:p></p>
<p class="MsoNormal">There were many false starts and iteration cycles to achieve
all of this, none of which is covered
here. The result is summarized in the
unavoidably-busy figure, <b>Figure 7</b>
just below, which includes a weight
statement that also displays mass fractions. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEib1iYYGsXpV0Q6ons-uxJKhlsmR2ZutXwwVtA4kpCIahyphenhyphenZ1FTgiHoZliffht5wTzusrNavgSQbgD97D5kCyQnVwZpFNWgftIFg2cMDwE5mEEEEvVp5bRMVkP7BxtMYs_tAobjfVmOpJ2oGIj5RrX_uhNmYrEr1SVeg6wX4aPGkFJB0CWy-nvFUwoyy9lpU/s1017/f12%20thrusters%20corrected%20r.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="1017" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEib1iYYGsXpV0Q6ons-uxJKhlsmR2ZutXwwVtA4kpCIahyphenhyphenZ1FTgiHoZliffht5wTzusrNavgSQbgD97D5kCyQnVwZpFNWgftIFg2cMDwE5mEEEEvVp5bRMVkP7BxtMYs_tAobjfVmOpJ2oGIj5RrX_uhNmYrEr1SVeg6wX4aPGkFJB0CWy-nvFUwoyy9lpU/w400-h240/f12%20thrusters%20corrected%20r.png" width="400" /></a></div><p class="MsoNormal">Figure 5 – Smaller Maneuvering Engines as Sized <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1WBbiN7QXN4Z41kOse_9B8XAH6MmWnPlz8PsJgnAZ_lZmn0QLZuzhkMDR9m6M1tZNMd_-xPri3k4YSOA8N1wW6Hz4k7PqGAc33iJL-J3JXPrn6pQNMlTb3f3_e2EbKgIa2_mMoDbkbhtqIyuzv9rFVWNM13MyhvlH7V-Lt_DDyBeKNc_qBy_Fy4B7uSxa/s1017/f13%20thruster%20feed%20options.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="610" data-original-width="1017" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1WBbiN7QXN4Z41kOse_9B8XAH6MmWnPlz8PsJgnAZ_lZmn0QLZuzhkMDR9m6M1tZNMd_-xPri3k4YSOA8N1wW6Hz4k7PqGAc33iJL-J3JXPrn6pQNMlTb3f3_e2EbKgIa2_mMoDbkbhtqIyuzv9rFVWNM13MyhvlH7V-Lt_DDyBeKNc_qBy_Fy4B7uSxa/w400-h240/f13%20thruster%20feed%20options.png" width="400" /></a></div><p class="MsoNormal">Figure 6 – Of the Options,
Pressurized Tanks Seemed Best <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiDeuiFWHIYsqyj6mFvJ_TqM0RknNI25ko8Zif5PKh-t0qzrBk_poiQeLpvcdXUBpZ2EosKOnXsIrM31Bn560KkDnmIwu_qTnEB42-Mot08keJhA055wzdSrEnNb5QXJ83KqEbt7xRu0By1v33XvZVuE-iIP8NPG7GKivKBR1Ep_NQbM2unKsENTglYLzmO/s998/f36%20final%20summary.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="681" data-original-width="998" height="272" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiDeuiFWHIYsqyj6mFvJ_TqM0RknNI25ko8Zif5PKh-t0qzrBk_poiQeLpvcdXUBpZ2EosKOnXsIrM31Bn560KkDnmIwu_qTnEB42-Mot08keJhA055wzdSrEnNb5QXJ83KqEbt7xRu0By1v33XvZVuE-iIP8NPG7GKivKBR1Ep_NQbM2unKsENTglYLzmO/w400-h272/f36%20final%20summary.png" width="400" /></a></div><p class="MsoNormal">Figure 7 – Summary Data for the Final Rough-Out Design <o:p></o:p></p>
<p class="MsoNormal">Of interest would be the various tank volumes. Bear in mind these are fully filled for the
mission to low circular Mars orbit at only low inclinations eastward, and also fully-filled for the long-range
suborbital mission (at low or high inclinations). The headers and maneuvering tanks are always
fully-filled, but the mains are only
partially filled for the shorter-range suborbital missions, so that entry mass is not too big.<o:p></o:p></p>
<p class="MsoNormal">Suborbital ranges from 9400 to just under 500 km were
examined in this study. Their entry
angles turned out to be a strong function of the suborbital mission
ranges. All of those suborbital entry angles
were considerably steeper than the return from the orbital mission. They were determined by feasible altitudes at
end-of-hypersonics, and by feasible peak
entry gee values.<o:p></o:p></p>
<p class="MsoNormal"><b>Figure 8</b> just below shows the final plots I got of
various flight data during entry and final descent. The Suborbital trajectories form trends, and the orbital data fall way off those
trends. Upper left is end-of-hypersonics
altitude and entry angle vs entry speed.
Upper right are the peak heating values during entry. Lower left are the trends of peak entry
gee, and average gee during the final
rocket-braked landings. There is a
numbered key relating the missions to each data point in each plot. No gee level exceeds 4.5, and no stagnation heating level exceeds 12.5
W/cm<sup>2</sup>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhg9KG5BZTkbomc_SZ9WQ1Ck-1diayz_Cz6rjIlzn_cM98K7D3ZpGWYVzbxAIrY0ffOqr6KISLUlpB78r1Fi7qYQeUL1EKuF9hzIWOx3pAXRg0A-Cva7vBfVCohjb6YqjJDNRdWVaACXzXGbFvW6m7yYklsOG5lyo84inndoP4eBGNS2MwJa4ty6VL2RzDb/s1091/f33%20basic%20EDL%20data.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="614" data-original-width="1091" height="225" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhg9KG5BZTkbomc_SZ9WQ1Ck-1diayz_Cz6rjIlzn_cM98K7D3ZpGWYVzbxAIrY0ffOqr6KISLUlpB78r1Fi7qYQeUL1EKuF9hzIWOx3pAXRg0A-Cva7vBfVCohjb6YqjJDNRdWVaACXzXGbFvW6m7yYklsOG5lyo84inndoP4eBGNS2MwJa4ty6VL2RzDb/w400-h225/f33%20basic%20EDL%20data.png" width="400" /></a></div><p class="MsoNormal">Figure 8 – Data for Entry,
Descent, and Landing (E, D, &
L) <o:p></o:p></p>
<p class="MsoNormal">Once again in <b>Figure 9</b> just below, the suborbital data for surface temperatures
also form trends versus entry speed,
with the orbital data falling far off of those trend lines, plus a numbered mission key. There are trends for surface temperatures at
the stagnation line, temperatures for
its lateral surfaces where flow is still attached, and temperatures for leeward separated wake
zone surfaces. These were figured for
thermal re-radiation exactly balancing convective-only input, as figured for a “dark” highly-emissive
surface, of thermal emissivity 0.8 as
representative. <o:p></o:p></p>
<p class="MsoNormal">Note that with the exception of only the longest-range
suborbital mission, all the rest of
these data are under 1600 F, and would
permit exposed-metal construction of 316L stainless steel, or of Inconel X-750, or something in that same class! <b><i>And that even includes the return from
the orbital mission!</i></b> Because of
the stagnation zone temperature approaching 2000 F on the longest-range
suborbital mission, there needs to be
some minimal heat protection in and near the stagnation zone. <o:p></o:p></p>
<p class="MsoNormal">In <b>Figure 10</b> just below, the format for surface pressures is
similar: trends of suborbital surface
pressure vs entry speed at stagnation,
at lateral sides, and in the
separated leeside wake. The orbital data
again fall far off the trend lines.
There is a numbered key to relate missions to individual data
points. Note that no mission, not even the longest-range suborbital, exceeds 0.19 atmosphere anywhere. That would be about 2.79 psi, very modest indeed. <o:p></o:p></p>
<p class="MsoNormal">Given the hot material strengths reported as part of <b>Figure
9</b>, that means even a fragile
extreme-low-density alumino-silicate ceramic composite could serve as heat
protection. So could ceramic fabric
blanket or quilt-type materials, if they
survive wind shear. Even a thin sheet
of 2000 F-capable metal overlying mineral wool insulation would serve, mounted only locally near the stagnation
line. <o:p></o:p></p>
<p class="MsoNormal"><b><u><span style="background: yellow; mso-highlight: yellow;">Conclusion: the “hopper” could easily be designed to also
serve as a low orbit taxi!</span><o:p></o:p></u></b></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjelaY9wZao4ETkMZzo2LOVhHRqzbi3DQzRBY7Qd8rZcvmadEY-6k46rccDeCybrPBsh78Jt-D3kn8EmTouenSxt7E53DoGK0x9mknOIq9PSL26XYqNY8iw7BX2JqIVmfhRs9V586eZdtvIQ00BVosN_0C2-ovHpJfk4EnYB-XF3XZOtgbZG28KK0vMW8-7/s1011/f35%20peak%20entry%20surface%20temps.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="1011" height="233" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjelaY9wZao4ETkMZzo2LOVhHRqzbi3DQzRBY7Qd8rZcvmadEY-6k46rccDeCybrPBsh78Jt-D3kn8EmTouenSxt7E53DoGK0x9mknOIq9PSL26XYqNY8iw7BX2JqIVmfhRs9V586eZdtvIQ00BVosN_0C2-ovHpJfk4EnYB-XF3XZOtgbZG28KK0vMW8-7/w400-h233/f35%20peak%20entry%20surface%20temps.png" width="400" /></a></div><p class="MsoNormal">Figure 9 – Trends of Surface Temperatures vs Entry Speed <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjQ6lOkgsCIK5lkZM_AxMwmC5v9jciX0qMeB4MKyKb9vkIx6xpNhJw8_btcfvrezpKyhgCyc32mdu97cCubMZzPConc5eoeRF6KlF1nvAQUnldik5dZFT5bHWoHDcUWJa9jYlLu8YS8pC104SGt3t6tQRIVUczrDN2kY6P3zwAM0szaCXfnncyUk322X6ig/s1011/f34%20peak%20entry%20surface%20pressures.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="1011" height="233" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjQ6lOkgsCIK5lkZM_AxMwmC5v9jciX0qMeB4MKyKb9vkIx6xpNhJw8_btcfvrezpKyhgCyc32mdu97cCubMZzPConc5eoeRF6KlF1nvAQUnldik5dZFT5bHWoHDcUWJa9jYlLu8YS8pC104SGt3t6tQRIVUczrDN2kY6P3zwAM0szaCXfnncyUk322X6ig/w400-h233/f34%20peak%20entry%20surface%20pressures.png" width="400" /></a></div><p class="MsoNormal">Figure 10 – Trends of Surface Pressures vs Entry Speed <o:p></o:p></p><p><b><u>Update 11-22-2023</u>:</b> The following Figure 11 illustrates exactly why the surface emissivity must be high (a very dark or black surface color, with a dull finish). There are exposed metallic construction solutions and a refractory solution with simple alumino-silicate ceramics, especially away from the stagnation zone, if emissivity is high. There ae only ablative solutions available if emissivity is low.</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiFG93H2TsClxA4lR-mmYGiXV664hyphenhyphenVo7LZH5R9a2j0w0hET3oZmpdxyfOLSETaK1eL4QBOYDI4ol0ILWFRLUijTvJgOIH4o5OsmrD5Qdo0KVDfu2HjkcK4NGj4GtYCEZjpjc8NhQGNGZA6s6sR1y_xjYdB8ahx-cBX9rpwo5EyI0jPT9fr5JWgGN_g_8rW/s1003/hopper%20taxi%20heating.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="668" data-original-width="1003" height="266" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiFG93H2TsClxA4lR-mmYGiXV664hyphenhyphenVo7LZH5R9a2j0w0hET3oZmpdxyfOLSETaK1eL4QBOYDI4ol0ILWFRLUijTvJgOIH4o5OsmrD5Qdo0KVDfu2HjkcK4NGj4GtYCEZjpjc8NhQGNGZA6s6sR1y_xjYdB8ahx-cBX9rpwo5EyI0jPT9fr5JWgGN_g_8rW/w400-h266/hopper%20taxi%20heating.png" width="400" /></a></div>Figure 11 -- More Detailed Hopper/Taxi Heating Data<br /><p><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-63903162313184481952023-11-04T10:56:00.002-05:002023-11-04T10:56:46.719-05:00Surface Freight Transport on Mars<p>I have seen many notions for surface transport discussed on
the New Mars forums, mostly in the
planetary transportation topic. <b><i>Nearly
all of these suffer from the very serious downside of requiring the emplacement
of significant infrastructure on Mars</i></b>,
something very effort-intensive and very expensive, here on Earth. With transport costs to Mars being
“astronomical” for the foreseeable future,
that is a fatal requirement for anything we might consider on Mars.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">What one really needs to do first is determine a very
fundamental characteristic:<span style="mso-spacerun: yes;"> </span>is the cargo
time-sensitive,<span style="mso-spacerun: yes;"> </span>or is it not?<span style="mso-spacerun: yes;"> </span>Transport of people over long distances is a
very time-sensitive problem,<span style="mso-spacerun: yes;"> </span>for which
the solution is some sort of flight,<span style="mso-spacerun: yes;"> </span><b>as
Figure 1 suggests</b>.<span style="mso-spacerun: yes;"> </span>Most bulk freight
is not time-sensitive at all,<span style="mso-spacerun: yes;"> </span>and can
travel very slowly,<span style="mso-spacerun: yes;"> </span>similar to scheduled
rail freight here on Earth.<span style="mso-spacerun: yes;"> </span><b><i>Considering
the expensive infrastructure issue,<span style="mso-spacerun: yes;"> </span>what
we need is a train that does not need any tracks.<span style="mso-spacerun: yes;"> </span><o:p></o:p></i></b></p>
<p class="MsoNormal">That small portion of freight to be transported that is time-sensitive
is likely to be medicines and similar:<span style="mso-spacerun: yes;">
</span>small packages that can fly with the people. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjqdYZKCJnabwsE0gTKPyjZ4iSolEXIfNK9m7bRonSuT2YtLMOfzCnoyoM5zbCoS5cWVk7tIKU_W9WnVtfAVAB7i78hcXPEY9ilEugIPAKA64UTuxQR0mj0vz0h-FCynF4aLUBNkjHEZveJ5eiIE8YQTfpCWGWIp59mf9Tn5Ybw3LpuS3wqHokJGpU7ar_J/s1009/f1%20why%20truck%20train.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjqdYZKCJnabwsE0gTKPyjZ4iSolEXIfNK9m7bRonSuT2YtLMOfzCnoyoM5zbCoS5cWVk7tIKU_W9WnVtfAVAB7i78hcXPEY9ilEugIPAKA64UTuxQR0mj0vz0h-FCynF4aLUBNkjHEZveJ5eiIE8YQTfpCWGWIp59mf9Tn5Ybw3LpuS3wqHokJGpU7ar_J/w400-h246/f1%20why%20truck%20train.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Time-Insensitive Freight Should Go Slow, On The Surface<o:p></o:p></p>
<p class="MsoNormal"><b><i>The critical thing here is to minimize the
infrastructure we must emplace to get any of these transportation ideas going.</i></b> The more you have to build, the more it is going to cost, for both the efforts, and the materials, plus the equipment with which to build
it. That’s true here on Earth, which is why many of these ideas have never
really been widely built, even
here. On Mars, it will cost a lot more still, because of the interplanetary transport costs, plus the development costs of re-making the
materials and processes into those that will actually work on Mars. As <b>indicated in Figure 2</b>, about the only thing we have available that
would actually work on Mars, too, would be graded dirt roads, with “truck trains” moving on them. True rail would cost less to run, being lower friction, but we have no way to manufacture steels on
Mars, including the extreme-cold adapted
steels necessary there, and no way to
make the ties, whether from steel, concrete,
or wood.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjBVzJyKhW6-YeaqzpyOLzzDXd_Yu4G88vRU7Z60rCP8-ajHQlhRO-tRcx7jva37jUY_bnKQhXpROX7QZrt0ptamW15v5Kvio2dJicVaBHXLLkq_4oOaZQYuwn6Njkw3622Y7yaEqSK3BkfDsH2Mbg5Kt8QQNnIarp_wESlMLk9Dw3MJJVA1q4SGNOh6YhT/s1009/f2%20infrastructure.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjBVzJyKhW6-YeaqzpyOLzzDXd_Yu4G88vRU7Z60rCP8-ajHQlhRO-tRcx7jva37jUY_bnKQhXpROX7QZrt0ptamW15v5Kvio2dJicVaBHXLLkq_4oOaZQYuwn6Njkw3622Y7yaEqSK3BkfDsH2Mbg5Kt8QQNnIarp_wESlMLk9Dw3MJJVA1q4SGNOh6YhT/w400-h246/f2%20infrastructure.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – The Real Requirement is That There Be Almost No
Infrastructure to Emplace<o:p></o:p></p>
<p class="MsoNormal"><b><i>So, it is the
slow-moving “truck train” hauling the time-insensitive freight on graded dirt
roads that we need for Mars. </i></b>This
is really just a tractor pulling a string of freight wagons, but internal combustion engines as we know
them here on Earth are “right out”. You
not only have to carry the fuel, but
also the oxygen with which to burn it,
which considerably outweighs the fuel.
And unless you want to <u>completely redesign the engines</u> to handle
500-1000 C higher flame temperatures,
you will also have to carry diluent gas,
which in turn far outweighs the oxygen.
<b><i>So, in any practical
sense, you are looking at
electric-powered tractors. <o:p></o:p></i></b></p>
<p class="MsoNormal">These can be <u>solar</u> electric powered, however.
On a sunny day here on Earth at lower latitudes, there is roughly a horsepower’s worth of
energy per square yard falling on the ground.
At Mars, there is about half
that. Call it 300 W per square meter as
a rule-of-thumb. Solar electric
conversion efficiencies are around 20%,
so about 60 W per square meter of collector surface could be had, for much of the day. The trick is then to fit the freight wagons
with solar panels on their roofs, all connected
electrically to power the tractor. You
just go very slow to stay within the power your solar collectors can
generate. But, most bulk freight is very
time-insensitive, so who cares?<o:p></o:p></p>
<p class="MsoNormal">This thing doesn’t need a crew, it can be self-driving between the spoil
berms created from grading the road. If
programmed with a keep-right feature,
traffic on these roads can be two-way.
The basic characteristics and features are <b>summarized in Figures 3
and 4</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh-wlWgMYiZ3eLyHRIp7bkkj07RB_GQwDAzfClYLicQCteUK_CXmBwrUqZOKDQ9TMIZVB6f5Z_h4FU9Z7ZozSuBpeR7gXsWQlcrWYEulTe6EIXuIe8QSjyh0yREchS3t8EscM94atNKfTmqOQrp62cOETPKQf8aMnvD3DrQ0S8V5rVq8vHjmxI4oGZHwZxU/s1009/f3%20truck%20train%20transport.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh-wlWgMYiZ3eLyHRIp7bkkj07RB_GQwDAzfClYLicQCteUK_CXmBwrUqZOKDQ9TMIZVB6f5Z_h4FU9Z7ZozSuBpeR7gXsWQlcrWYEulTe6EIXuIe8QSjyh0yREchS3t8EscM94atNKfTmqOQrp62cOETPKQf8aMnvD3DrQ0S8V5rVq8vHjmxI4oGZHwZxU/w400-h246/f3%20truck%20train%20transport.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – The “Truck Train” Concept<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiBvGqBvVl0J3VNmrZrSDfCbSfPfci_ZUxePL80l4lgwOEsL-Gsbp7gvgKGRTatKpe8kJWsGbDsX6p3sAc_l5M15Gvahm06ufVcSMogxWIGaWu7x9GeOdJHDdvrRtfQs-idtF2dsNx2eu78s0Vo3n8bCflNfRf6MjNZLab4h-TYXbqqNMKtv-s71CywuR4g/s1009/f4%20roadway%20and%20control.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiBvGqBvVl0J3VNmrZrSDfCbSfPfci_ZUxePL80l4lgwOEsL-Gsbp7gvgKGRTatKpe8kJWsGbDsX6p3sAc_l5M15Gvahm06ufVcSMogxWIGaWu7x9GeOdJHDdvrRtfQs-idtF2dsNx2eu78s0Vo3n8bCflNfRf6MjNZLab4h-TYXbqqNMKtv-s71CywuR4g/w400-h246/f4%20roadway%20and%20control.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – This Is How It Is Controlled<o:p></o:p></p>
<p class="MsoNormal"><b><i>Now we need to verify feasibility with some numbers.</i></b> Trucks and train cars here on Earth are about
1/3 structure and 2/3 payload, sometimes
¼-3/4. Call it 30-70, as a rule of thumb. For a 100 metric ton loaded freight
wagon, you are looking at 70 tons of
payload, and 30 tons of chassis, wheels,
drawbars, containing-structure
for the freight, and solar panels on the
roof. <o:p></o:p></p>
<p class="MsoNormal">These things could use the very same giant rubber tires we
use for off-road construction and mining work here on Earth, but we may need to heat them slightly, to prevent their cracking in the cold on
Mars. That can be done. Tires on relatively smooth, firm dirt have crudely 10 times the friction
coefficient of tires on a paved road.
Using the Mars weight of a 100-ton loaded freight wagon, augmented for climbing a 10% grade, the drawbar power (drag x speed) is 20 KW at
0.33 m/s, and quite a bit higher as you
go faster. There’s room for at least around
300-ish square meters of collectors atop each wagon. Which in turn is why you select the lower of
the speeds <b>given in Figure 5</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgBRSyzWWaJRylFOnopkWd3oBaMtTB-ZqlWMQkuozf2ekjy8IH-0rAGn0WluUSBdZ24sNWmL45KodcwAy1oZehwAPApPTqI8kiVLBw4EKyo-141f3X4y1Wpas21rfmb7ranfI7Z_ZXsDYMJPwZZm8AJlg8KyYw2XLW72f6yzg_pdfu4hAUyQ9z-3ZZRheAK/s1009/f5%20tractor%20power.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgBRSyzWWaJRylFOnopkWd3oBaMtTB-ZqlWMQkuozf2ekjy8IH-0rAGn0WluUSBdZ24sNWmL45KodcwAy1oZehwAPApPTqI8kiVLBw4EKyo-141f3X4y1Wpas21rfmb7ranfI7Z_ZXsDYMJPwZZm8AJlg8KyYw2XLW72f6yzg_pdfu4hAUyQ9z-3ZZRheAK/w400-h246/f5%20tractor%20power.png" width="400" /></a></div><p class="MsoNormal">Figure 5 – This Is Why It Works (By the Numbers)<o:p></o:p></p>
<p class="MsoNormal">You can pull many freight wagons with a big powerful
tractor. Numbers are given in the figure
for 10 and 100 wagons. Power at the
tractor is between 0.2 to 2 MW for these numbers. That’s about like the power of a big mine
loader as currently built here on Earth (somewhere near 1000 HP). Which in turn means the same electric tractor
that can pull the “truck train” can also be the road grader! You just power it
differently. <o:p></o:p></p>
<p class="MsoNormal">The kind of thing I have in mind as the basis for designing
the Mars tractor is <b>pictured in Figure 6</b>. That is a big mine loader vehicle. These are very tough, very powerful machines. That is the very thing we need for doing road
grading, and pulling heavy trains of
freight wagons. We just have to make it
work on Mars, in the cold and the
near-vacuum. That is why electric
propulsion is preferred, and a
pressurized operator’s cabin is required. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj9kSeh5lprjSt3O5IGKNIaQI9o-8aJ3JZT29TkTTToyWdlhyphenhyphenxZFx3WWo7NE4HN_JZqQ166Iodsx0i3HmHv_hNM_6mDNeo5v291Wks4A9HJ3nJu7Eo547DdbjZMD8rIIvdqsuDATAHwz4BPunYaswUSDDN-BfId1jfUQ9c-ac7_3GSAVjv3sx-5ldZ3J6er/s712/f6%20mine%20loader%20jpeg.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="534" data-original-width="712" height="300" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj9kSeh5lprjSt3O5IGKNIaQI9o-8aJ3JZT29TkTTToyWdlhyphenhyphenxZFx3WWo7NE4HN_JZqQ166Iodsx0i3HmHv_hNM_6mDNeo5v291Wks4A9HJ3nJu7Eo547DdbjZMD8rIIvdqsuDATAHwz4BPunYaswUSDDN-BfId1jfUQ9c-ac7_3GSAVjv3sx-5ldZ3J6er/w400-h300/f6%20mine%20loader%20jpeg.jpg" width="400" /></a></div><p class="MsoNormal">Figure 6 – The Mars Tractor Is a Known Revision of Something
Like This That We Already Build</p><p class="MsoNormal"><o:p></o:p></p><p class="MsoNormal">The basic notion is <b>illustrated in Figure 7</b>. Something similar to a big mine loader is the
design basis. You remove the dump hopper
and replace it with a battery bank and solar panels. You remove the diesel power plant and replace
it with the appropriate electric drive motors and gearing. You
replace the operator cab with a pressurized operator’s cabin for use on
Mars. <o:p></o:p></p><p class="MsoNormal">Rigged with a blade,
and powered by a nuclear generator aboard a shielded trailer, you operate it manned for grading the
road. Unmanned and self-driving, you pull the “truck train” with it, powered by the solar panels on the freight
wagons.<o:p></o:p></p><p class="MsoNormal"><b><i>It will have to be manned for grading the road.</i></b> Even here on Earth, “dirt work” is something that has so far
proven to be not-automatable, or it already
would have been automated. That’s slow
multi-pass work, with the operator’s
judgement and skill showing up, as
knowing when the work is done “right” by the appearance of what he has
done. This aspect will be the same on
Mars as it is here.<o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal"><b><i>It does not have to be manned to pull the “truck
train” on the finished roadway. </i></b> You remove the blade, and delete the nuclear power trailer. However,
I would leave the pressurized cabin in place and rigged with
supplies, for the off-chance unforeseen
event that would require a human driver. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi5f51TdapQMGCOFlHAwFF58076yDnuBnLGOx_oG6HC8uo0Ux4cRdk5USw-PPAgeOFY8347u1bS_23-ccdIAPH-MZj3Yjlaujemfuy6T_0cC9lHQ5yNPkd-V1qGwU0KIJaHa_L6s815GsHDr1gqD11bqlHDOg6DrQKPi2AVm9eZwTgii3hbDGql9qRgzwgo/s1009/f7%20modified%20mine%20loader.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi5f51TdapQMGCOFlHAwFF58076yDnuBnLGOx_oG6HC8uo0Ux4cRdk5USw-PPAgeOFY8347u1bS_23-ccdIAPH-MZj3Yjlaujemfuy6T_0cC9lHQ5yNPkd-V1qGwU0KIJaHa_L6s815GsHDr1gqD11bqlHDOg6DrQKPi2AVm9eZwTgii3hbDGql9qRgzwgo/w400-h246/f7%20modified%20mine%20loader.png" width="400" /></a></div><p class="MsoNormal">Figure 7 – Converting a Mine Loader Design into a Mars
Tractor Design That Does Two Jobs<o:p></o:p></p>
<p class="MsoNormal"><u>The only real trouble is shipping such large pieces of
equipment to Mars</u>. These will have
to be shipped as individual components,
and assembled there on the planet Mars. That will be true until some
very large interplanetary vehicles indeed,
eventually become available. But
it can be done, with what we know and
have available, right now!<o:p></o:p></p><p class="MsoNormal"><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com1tag:blogger.com,1999:blog-2675974463524895416.post-54448368495743470132023-11-01T14:10:00.001-05:002023-11-01T14:10:47.693-05:00Rocket Hopper for Mars Planetary Transportation<p class="MsoNormal">The future presence of bases and settlements upon Mars
brings the need for transportation of freight and people about the planet. A little thought reveals the two
categories, freight and people, are fundamentally different in their
requirements. Most freight is not
time-sensitive, while people are. <o:p></o:p></p>
<p class="MsoNormal">Freight not time-sensitive needs to go by slow surface
transportation, without emplacing
expensive and effort-intensive infrastructure to make that possible. Such infrastructure is often very expensive
here, and the costs there are likely to
be quite catastrophic. Freight needs
something similar to rail transport here,
but without the tracks. A robot
“truck train” on a graded dirt road is the answer.<o:p></o:p></p>
<p class="MsoNormal">People are time-sensitive,
and need to fly in order to cover large distances rapidly. But physics in an atmosphere so thin, that it is first cousin to the vacuum of
space, rules out air transport by
airplanes or helicopters as we know them on Earth. It also rules out any form of
“lighter-than-air” transport, since
buoyant lift forces are proportional to differences in densities, and those are so vanishingly-low on
Mars. That pretty much leaves suborbital
rocket travel (the “rocket hopper”), the
topic of this document. The same could
be used for that tiny portion of freight that really is time-sensitive. <o:p></o:p></p>
<p class="MsoNormal"><b>Suborbital Trajectories<o:p></o:p></b></p>
<p class="MsoNormal"><o:p></o:p></p><p>I used my “orbit basics” spreadsheet, “R V q orbits” worksheet, to model suborbital trajectories about the
planet, using ellipses that lie mostly within
the planet. It was necessary to sharply
limit the trajectory angle below local horizontal at entry interface altitude
(135 km for Mars), in order to limit
peak entry deceleration gees and peak heating,
and to achieve altitudes from which rocket-braked landings were feasible
in terms of timelines and deceleration gees.
The results I was able to obtain are <b>shown in Figure 1</b>. The longest-range of these trajectories is
very nearly antipodal (10,637 km). </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjlNFC1M-LIbTCjBKK_YTpFk5sL8NLkb19h4NRBfpsFAl9I9K7FL3GzOF_Q82L7XRV_hZYatzXCrdMNWbvfMxooZMfyVgQvHkvLcADjN1I7bO24iawy3TjZ2GrZMQlP_P8ye2kgJr9d40vmT21QUHsCpMyFYaoxWwHpsmEh8Fn0vQo8psnmuxl651bsW1RD/s1009/suborbit%20traj.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1009" height="246" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjlNFC1M-LIbTCjBKK_YTpFk5sL8NLkb19h4NRBfpsFAl9I9K7FL3GzOF_Q82L7XRV_hZYatzXCrdMNWbvfMxooZMfyVgQvHkvLcADjN1I7bO24iawy3TjZ2GrZMQlP_P8ye2kgJr9d40vmT21QUHsCpMyFYaoxWwHpsmEh8Fn0vQo8psnmuxl651bsW1RD/w400-h246/suborbit%20traj.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Suborbital Trajectories<o:p></o:p></p>
<p class="MsoNormal"><b>Entry,
Descent, and Landing<o:p></o:p></b></p>
<p class="MsoNormal">I used the “entry spreadsheet” spreadsheet file, worksheet “Mars variations”, to compute estimates for the worst case
(highest entry speed) entry trajectory,
using the old 1953-vintage analysis used for warheads by H. Julian Allen
and his colleagues. This stuff was
declassified and taught in engineering schools in the late 1960’s. It is not the most accurate thing to use, but it gets you “well into the ballpark” with
very simple, essentially-by-hand, calculations.
These are easily put into a spreadsheet.<o:p></o:p></p>
<p class="MsoNormal">The highest speed at entry is the most challenging, since all the entry angles are about 15
degrees, so that analysis is <b>depicted
in Figure 2</b>. It was crucial to get
the hypersonic ballistic coefficient of the hopper vehicle down under 200
kg/sq.m, and preferably under 180
kg/sq.m. It proved feasible with the
final design to get well under that requirement with the value shown in the figure, although most design approaches will fail in
that respect! Getting the heating down
to something easily handled required the largest-possible “nose radius”. The value shown was dictated by the design
approach taken here, and is quite easily
handled, although it is too much for
metal exposed at the stagnation zone. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiOPsxxZ6_kdITYOnPpJNms5WgVnZfue8DU1RG1sX-KZiLsnXIa_dzknwa34SAFOxkgEUcVNobxrXt6l14RpetOiWO7Kue1Tvn9yYAWMwnXH4UAHzDMFyYL8X7K1O3nQTXlgNIhawvaciW6_DakdgPAfODvL8HzynVY57Bzm2ajKvFGJf7IVEIX74PuUUzj/s1016/final%20size%20part%203.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="623" data-original-width="1016" height="245" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiOPsxxZ6_kdITYOnPpJNms5WgVnZfue8DU1RG1sX-KZiLsnXIa_dzknwa34SAFOxkgEUcVNobxrXt6l14RpetOiWO7Kue1Tvn9yYAWMwnXH4UAHzDMFyYL8X7K1O3nQTXlgNIhawvaciW6_DakdgPAfODvL8HzynVY57Bzm2ajKvFGJf7IVEIX74PuUUzj/w400-h245/final%20size%20part%203.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Estimates for Entry Conditions<o:p></o:p></p>
<p class="MsoNormal">The hypersonic aerobraking is largely over once the vehicle
has decelerated to about local Mach 3,
which is roughly 0.7 km/s speed in the Martian atmosphere. Slower than Mach 3, a blunt object is no longer hypersonic, and the assumption of a constant ballistic
coefficient fails. It would be folly to
continue the hypersonic estimate past that point, for that very reason. <o:p></o:p></p>
<p class="MsoNormal">Not included in the approximate analysis is the effect of
trajectory “droop” to steeper angles due to gravity. That mostly happens after the peak
deceleration point, which is actually
rather close to the end-of-hypersonics (Mach 3) point. Peak heating occurred slightly earlier. For purposes of “reasonable
approximation”, I just used the
hypersonic endpoint altitude as obtained,
but I presumed the trajectory was headed about 45 degrees downward at
the Mach 3 point. <o:p></o:p></p>
<p class="MsoNormal">Ignoring the effect of the potential energy associated with
the Mach 3 point altitude, it is the
same Mach 3 point speed of 0.7 km/s that we have to “kill” with last-second rocket
braking, regardless of that
altitude. Appropriately factored for
losses and maneuvering, that is the
rocket braking delta-vee (dV) that we must have. I used factor 2, and would never use less than 1.5 under any
circumstances, in order to account for
losses, plus hover and divert
requirements. The Mach 3 point altitude essentially
determines how much time we have left before surface impact, which sets the required average gees, and thus the thrust required for any
particular vehicle mass. <o:p></o:p></p>
<p class="MsoNormal">I used simple high school-level geometry and physics/kinematics
to establish the average deceleration gees during the rocket-braking
landing. This is a very simple hand
calculation, <b>illustrated in Figure 3</b>. The 6 km altitude translates to an 8.5 km
path length down a straight line at 45 degrees.
At an undecelerated constant 700 m/s,
we are about 12.1 sec from impact,
as shown. Thus there is far too
little time available to deploy a chute,
much less expect any deceleration from it. <o:p></o:p></p>
<p class="MsoNormal">If we rocket-brake decelerate to zero, the average velocity down the path is only
350 m/s, and we have 24.3 sec to
touchdown, as shown. The change in speed is the 700 m/s. The change in time is the 24.3 sec. Their ratio is the average deceleration
required, which is 28.8 m/sec<sup>2</sup>, or some 2.94 standard gees.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgLwtr7gf-xAt_tJ2McXAH1COt_w8Dj2Tnas94rX28tqZoqi7i8wgFYMn1IevIdWxarBIQn9VPUXV1RY5SMjNEiGvGi_e8FlDseEVYyGekaateYWw6NPxzm0fsVfNuNN0IBQKILEHbDg9aB9RshvkWVNBpPL7gcoPQCrZJf6stJyFpV91GWn6O2HOl9QtZ3/s1016/final%20size%20part%204.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="623" data-original-width="1016" height="245" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgLwtr7gf-xAt_tJ2McXAH1COt_w8Dj2Tnas94rX28tqZoqi7i8wgFYMn1IevIdWxarBIQn9VPUXV1RY5SMjNEiGvGi_e8FlDseEVYyGekaateYWw6NPxzm0fsVfNuNN0IBQKILEHbDg9aB9RshvkWVNBpPL7gcoPQCrZJf6stJyFpV91GWn6O2HOl9QtZ3/w400-h245/final%20size%20part%204.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – Hand Calculations for the Rocket-Braked Landing<o:p></o:p></p>
<p class="MsoNormal">Therefore, we are
looking at roughly a 3-gee rocket-braked landing, with the gees felt over an interval 24-25 sec
long. That is easily handled by persons
not trained in resisting gees, if seated, even more so if reclined. Roller coaster riders endure worse all the
time. <o:p></o:p></p>
<p class="MsoNormal">The hypersonic aerobrake peak gees fall in the 6.6 to 6.7
range, but again for a short interval above
5 gees that is only around 20-25 sec long,
as well. That is more
difficult, but it is endurable, even by untrained persons as long as they are
fully physically fit, as also
experienced by some roller coaster riders.
It is anticipated that passengers will be riding while wearing some sort
of pressure suits. It would help if
these pressure suits had “gee-suit” features as well. Otherwise,
some passengers might temporarily faint,
if sitting up. <u>Any crew must
be trained to endure such gees</u>, <u>and
they must be wearing suits with “gee-suit” features</u>.<o:p></o:p></p>
<p class="MsoNormal"><b>“Rocket Hopper” Vehicle Design Concept and Estimates<o:p></o:p></b></p>
<p class="MsoNormal">I looked at 3 classes of possible design configurations
trying to meet the requirements of low hypersonic ballistic coefficient and
large “nose radius” simultaneously. Only
one approach satisfied those needs, and
still offered ways to mount landing legs for rough-field operations, plus a simple unobstructed engine bay. That was the cylindrical stack <b>depicted in
Figure 4</b>, but flown dead-broadside
to the oncoming stream during entry! <o:p></o:p></p>
<p class="MsoNormal">If the cylinder L/D ratio falls in the 4-to-6 range, that is enough larger blockage area to
greatly-reduce the hypersonic ballistic coefficient, despite the low hypersonic drag coefficient
of the cylinder shape. That shape was
required to keep tank construction lightweight.<o:p></o:p></p>
<p class="MsoNormal">As for the tanks,
these are main tanks that are integral components of the vehicle
airframe, but they also contain header
tanks. As the numbers worked out, about 15% of the tank volume is in the
headers, for course correction and
rocket-braking, with the other 85% in
the main tank volume for launch. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjK0eKSVZ6SfUTwKFTvZPp3ZgR1CztHer7rlI1AiPLewPYZ7C2CU6M4CrVcz2b6N_7mXVq_ZFfNHZ_JU2PraEaj_IGIeC3Y7ONNuTz9YDAnwl6Zr6pHk4V-Ag8fUKEXmJj20NpSHAwbzVFcQfkUZShSoYhamscOSiqZNAwFH7q9Rt6je4l8jM61DBgP1r-x/s1016/final%20size%20part%205.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="623" data-original-width="1016" height="245" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjK0eKSVZ6SfUTwKFTvZPp3ZgR1CztHer7rlI1AiPLewPYZ7C2CU6M4CrVcz2b6N_7mXVq_ZFfNHZ_JU2PraEaj_IGIeC3Y7ONNuTz9YDAnwl6Zr6pHk4V-Ag8fUKEXmJj20NpSHAwbzVFcQfkUZShSoYhamscOSiqZNAwFH7q9Rt6je4l8jM61DBgP1r-x/w400-h245/final%20size%20part%205.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Sketch Layout and Characteristics of the “Rocket
Hopper” Design Concept<o:p></o:p></p>
<p class="MsoNormal">As shown, the
smallest item was the engine bay, and
the largest item the cabin in which people ride. The tanks are stacked in the middle, so that center-of-gravity travel is not very
large as the propellants burn off. The
figure shows the layout, a weight
statement, seating arrangements, the basic trajectory-related notions, and some entry heat protection numbers. <o:p></o:p></p>
<p class="MsoNormal">Most but not quite all of the surface of this craft could be
exposed metal construction, if something
like a 316L stainless steel or an Inconel X-750 is used. Only near the stagnation line on the windward
side is something more heat-resistant required.
That could be a strip of ceramic tiles of some kind. Even low-density alumino-silicate ceramics
could be used, if blackened for high
emissivity, as peak entry pressures are actually
quite low. <o:p></o:p></p>
<p class="MsoNormal">Peak entry pressures are easily rough-estimated by simple
hand calculations at the peak deceleration gee point in the hypersonic entry
trajectory. If you know the mass at
entry, the peak gees acting upon that
mass give you the peak force decelerating the vehicle. Dividing that force by the frontal blockage
area gives you the average pressure applied to that area. The peak is at the stagnation zone, crudely twice the average value. <o:p></o:p></p>
<p class="MsoNormal">To run some of these numbers, I did create a custom “rocket hopper”
spreadsheet. It has multiple
worksheets, of which two are relevant here. Worksheet “veh” is set up to make the
calculations <b>illustrated in Figure 5</b>.
Worksheet “sections” is set up to make the calculations <b>shown in
Figure 6</b>. <o:p></o:p></p>
<p class="MsoNormal">In worksheet “veh”,
user inputs are yellow,
significant outputs are blue, and
things requiring iteration or verification are green. (The same color code applies to all my
spreadsheets.) It works in terms of mass
ratio-effective dV, which is the
end-of-burn V multiplied by an appropriate factor to account for gravity and
drag losses (about 1.02 on Mars). <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjyecZpLouSMNQE0s1E-3ZBMua4OVw_wOIB6KWPM09Mh_p15U-1rl1WRbbz4zVYxLlchI77l9KODZKpngB8NOuYpY8mxvNwwn0KC08mp5ljyM9XLI-0zqzJvtMjScsekaZfJYjNXh6ZAuY7mxewmT2aldkf9G2QwVptlmvEnDWr1zGL3V8mSSlDeocmkHkp/s1361/final%20size%20part%201.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="769" data-original-width="1361" height="226" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjyecZpLouSMNQE0s1E-3ZBMua4OVw_wOIB6KWPM09Mh_p15U-1rl1WRbbz4zVYxLlchI77l9KODZKpngB8NOuYpY8mxvNwwn0KC08mp5ljyM9XLI-0zqzJvtMjScsekaZfJYjNXh6ZAuY7mxewmT2aldkf9G2QwVptlmvEnDWr1zGL3V8mSSlDeocmkHkp/w400-h226/final%20size%20part%201.png" width="400" /></a></div><p class="MsoNormal">Figure 5 – Numbers Run for the Design Concept, Part 1<o:p></o:p></p>
<p class="MsoNormal">The 3.6 km/s orbit speed at launch multiplied by that factor
is the 3.672 km/s dV shown. The course
correction budget is 8% of the orbit apoapsis speed, which is where corrections should be
made. For the highest-speed case, that is the 0.215 km/s shown. And at factor 2, the 1.4 km/s dV is what the rocket-braking
burn must be capable of, the same for
all cases. <o:p></o:p></p>
<p class="MsoNormal">I used a launch liftoff 1.5 gees as required for initiating
good ascent kinematics, same as here on
Earth. 0.1 gee for course corrections is
probably just a lower limit. I used 3.5
gees at landing to get some margin over the average 3 gees determined above. <o:p></o:p></p>
<p class="MsoNormal">The specific impulse input of 352 sec shown is justified by
the analysis given below. The sum of the
dV values is the total dV to be delivered for the mission, which sets vehicle mass ratio and propellant
mass fraction. 1 – propellant fraction –
inert fraction is the available payload fraction. <o:p></o:p></p>
<p class="MsoNormal">There is an “inert mass fraction build-out” block
shown. We can argue about the component
inputs, but their sum is likely “in the
ballpark” no matter exactly what inputs one uses. The same is true of the size payload
block. You are looking at the weights
for a person, his pressure suit, a few hours worth of oxygen, water,
and food, plus some luggage. It’ll be around 0.2-0.25 metric tons per
person, almost no matter what, when you sum it up. That and how many people
are aboard, sets the payload mass, which ultimately sets the weight statement. <o:p></o:p></p>
<p class="MsoNormal">The “heat shield re-radiation” block presumes convective
heating is balanced by re-radiation from hot exposed surface materials. It uses the peak heating from the entry
analysis, plus good guesses for surface
emissivity and the effective temperature of the surroundings that must receive
that re-radiated heat. A highly-emissive
surface is typically 0.8, while low is
0.2. 400 R for the surroundings is -60
F. <o:p></o:p></p>
<p class="MsoNormal">The “run weight statement and size thrusts” block does
exactly that. Payload mass divided by
payload fraction is ignition mass.
Ignition mass time inert fraction is inert mass, and times propellant fraction is propellant
mass. Payload plus inert is
burnout, and burnout plus propellant is
ignition. Each burn has a dV that sets
its mass ratio. That in turn sets start
and end-of-burn masses, the difference
being propellant used for that burn. The
sum of those propellant masses used must equal the total propellant mass
already figured. <o:p></o:p></p>
<p class="MsoNormal">There is a “check pressure on heat shield” block that uses
the hypersonic ballistic coefficient and the hypersonic drag coefficient as
inputs, plus the initial mass at
entry, and the peak entry gees. It computes the mass per unit blockage area
from the ballistic coefficient and drag coefficient, then the peak entry force from that and the
entry mass and peak entry gees. It then
divides force by blockage area for average pressure, and doubles that for the stagnation pressure
estimate, reported in a variety of units
of measure. <o:p></o:p></p>
<p class="MsoNormal">The ”heating other locations” block gets you equilibrium
temperatures in degrees F for stagnation,
“typical” lateral, and separated
wake zone locations. The stagnation peak
heating is reduced by a factor of 3 for “typical” lateral surfaces, and by a factor of 10 for separated wake zone
surfaces. <o:p></o:p></p>
<p class="MsoNormal">The only other block shows launch dV available as a function
of percent max propellant loaded on board.
The course correction and landing dV values are not changed. This would be useful trying to relate range
to propellant load required.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjMVDF_Er0A0cFeGYti0XVaI4FrLFGwh8pCtmiRyECivZxWjBbvDfpamb1fPnvZQAfe5TlnSspw5zrq0zck0sYBH2HsRGtTVgyJydnZPQmcrxYVaVY0mK5RNmvJwVfAXFHt-8ABjAtzY0PX3aOuXptt6W_CkmxQaAlxgy8f3O2-SzJkqvHnqYBu2YL8oKPP/s1521/final%20size%20part%202.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="625" data-original-width="1521" height="164" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjMVDF_Er0A0cFeGYti0XVaI4FrLFGwh8pCtmiRyECivZxWjBbvDfpamb1fPnvZQAfe5TlnSspw5zrq0zck0sYBH2HsRGtTVgyJydnZPQmcrxYVaVY0mK5RNmvJwVfAXFHt-8ABjAtzY0PX3aOuXptt6W_CkmxQaAlxgy8f3O2-SzJkqvHnqYBu2YL8oKPP/w400-h164/final%20size%20part%202.png" width="400" /></a></div><p class="MsoNormal">Figure 6 – Numbers Run for the Design Concept, Part 2<o:p></o:p></p>
<p class="MsoNormal">The “sections” worksheet works out the proportions of the
engine bay, tankage, and cabin sections of the ship, plus some other pertinent results. It needs the r-ratio (oxidizer/fuel fuel mass
ratio) for the engine, and the specific
gravities of the fuel and oxidizer materials.
It also needs as inputs the mass at start of entry (after course
corrections from the weight statement),
and a hypersonic drag coefficient in crossflow for the presumed vehicle
shape, which in this case is a circular
cylinder. <o:p></o:p></p>
<p class="MsoNormal">It also needs as inputs the propellant masses for each
burn. It works out from these the masses
of oxidizer and fuel for each burn, and
their volumes. This is done in an
untitled block top center of the page. <o:p></o:p></p>
<p class="MsoNormal">There is an “engine resize” block that takes the engine
characteristics modeled elsewhere, and
rescales them to the correct thrust size.
Those inputs are the modeled vacuum thrust, the required vacuum thrust, the throat and exit diameter sizes, and the effective average half-angle of the
supersonic expansion bell. <o:p></o:p></p>
<p class="MsoNormal">There is a “tanks figured on totals, with headers inside TBD” block. It has vehicle diameter and a sort of
interstage length between the tanks as inputs.
It works out all the lengths and L/D ratios, and requires the “right” vehicle diameter to
keep a full sphere as the smaller tank,
as well as to be consistent with the seating arrangements in the “cabin”
block. The seating is an input number per
level, and number of levels, consistent with the total number of
people. There has to be room for seating
at the input diameter in the “tanks” block,
and the input seat pitch is the spacing between levels. Note how the empty main tank shells protect
the propellants in the header tanks from the effects of entry heating.<o:p></o:p></p>
<p class="MsoNormal">The “engine bay” block takes the resized dimensions of the
engine (overall rough estimate of length,
and the exit diameter, and uses
these with inputs for number of engines and the spacing requirements for
gimballing, to estimate min dimensions
of the engine bay. Its diameter should
never exceed the input diameter in the “tanks” block. <o:p></o:p></p>
<p class="MsoNormal">From there, the
“overall vehicle” block puts together these results into estimates for the
length and L/D ratio (which ought to be in the 4 to 6 range) of the entire
vehicle stack, and with inputs for entry
mass and hypersonic drag coefficient,
estimates the broadside-entry ballistic coefficient. Too low an L/D will
get you too low a blockage area and too high a ballistic coefficient. Too high an L/D is a topple-over risk, or at least a risk of bigger, heavier landing legs. <o:p></o:p></p>
<p class="MsoNormal"><b>Creating the Re-Sizable Engine Data<o:p></o:p></b></p>
<p class="MsoNormal">I used the “rocket nozzle” spreadsheet file, worksheet “rocket noz”, to rough-size a suitable engine and calculate
a reliable performance estimate for it.
This was actually one of the first things I did. This engine burns liquid methane fuel with
liquid oxygen oxidizer, similar to
SpaceX’s Raptor and Blue Origin’s BE-4.
These propellants are thought to be manufacturable in situ on Mars. This entire design study assumes that to be
true.<o:p></o:p></p>
<p class="MsoNormal">I chose not to push the state of the art, given the troubles SpaceX had getting to
current Raptor-2 performance levels.
This engine is only sized for a chamber total pressure delivered to the
nozzle entrance of 2000 psia, and only a
3:1 pressure turndown ratio, although
this analysis does presume a full-flow cycle with no dumped bleed gas. The nozzle is assumed to be an 18-8 degree
curved bell, with a throat discharge
coefficient of 0.995. Both c* and r are
presumed functions Pc, using the data <b>shown
in Figure 7</b> as the inputs. Similar
data for other propellant combinations are in the “prop comb” worksheet. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhz_Iz1BGmcrgWhtACJK1eWEcDZQ5URHd65oQyge-QjYSVYyhzzgduf0ypmGosjGsHeFhLS1486ttP9YKe5H5QUZiF23e3RX_DyRm0HpByUyj0rqIcN2eFG_FTiF6zNSac93OgZ7b_9gwmC3haiV1n4GvPCPWStKarHL4ZIomZd_TEIXLNQLg-KgluF0v7o/s1761/engine%20sizing.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="629" data-original-width="1761" height="143" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhz_Iz1BGmcrgWhtACJK1eWEcDZQ5URHd65oQyge-QjYSVYyhzzgduf0ypmGosjGsHeFhLS1486ttP9YKe5H5QUZiF23e3RX_DyRm0HpByUyj0rqIcN2eFG_FTiF6zNSac93OgZ7b_9gwmC3haiV1n4GvPCPWStKarHL4ZIomZd_TEIXLNQLg-KgluF0v7o/w400-h143/engine%20sizing.png" width="400" /></a></div><p class="MsoNormal">Figure 7 – The Resizable Engine For This Study<o:p></o:p></p>
<p class="MsoNormal">The nozzle expansion was designed at full Pc = 2000
psia, expanded to 5.97 psia at its exit
plane. This gave us a nozzle right at
incipient separation when operated at 2/3 Pc, at sea level on Earth, allowing easy open-air testing on Earth. It is unseparated for that power setting (or
higher), but cannot be operated at lower
settings at sea level, because the
nozzle will separate. <o:p></o:p></p>
<p class="MsoNormal">Thrust was sized at an arbitrary 10,000 lb in vacuum at full
Pc. Performance in the near-vacuum of
Mars’s atmosphere is indistinguishable from true vacuum performance, so the vacuum data were used for this
study. There is very little specific
impulse variation from 1/3 to full thrust,
reflecting pretty much only the variation of c* with Pc. I used the 2/3
power value of 352 sec as “typical” of operation at any throttle setting. Thrust and flow rates scale with throat
area, dimensions scale with throat
diameter.<o:p></o:p></p>
<p class="MsoNormal"><b>Final Note<o:p></o:p></b></p>
<p class="MsoNormal">Most of the spreadsheets used here are part of the course
materials that I created for orbits and vehicle sizing.
Only the custom spreadsheet used for vehicle characteristics as calculated here, is not part
of those course materials. Those course materials are available on the New Mars forums, in the "interplanetary transportation" topic, "orbit mechanics class traditional" thread. <o:p></o:p></p><p><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-86534801870931152992023-10-23T11:14:00.000-05:002023-10-23T11:14:10.416-05:00Only Force Stops Bullies<p>Lost among all the war news from the middle east is more
bullying conducted by the Chinese. They
are actually ramming vessels belonging to the Philippines, according to news reports. That risks lives, and the longer this evil behavior goes
on, the more likely it is that people
will die. </p><p><br /></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEituin7PMfYpcGuEcXeZ8W7rVSLq1qsGsaQkFIRXLMrRDh2lwmOkyvZZtKkBHl1YTXUhXedzCVM5j_V-_qFj-mzqlx2reC4uHnmUVm7n5RblAoDI0o5wwlwfKDuK1M_OOoQY-Z7lxDI5ZOLOvWWy4vmsllQUsQgkm5sxCyZdJyNpZ26QWO-KuOcIbCsm0zV/s1200/2023-10-06T060942Z_867183703_RC2FL3AHA8TD_RTRMADP_3_SOUTHCHINASEA-PHILIPPINES-1200x800.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="800" data-original-width="1200" height="266" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEituin7PMfYpcGuEcXeZ8W7rVSLq1qsGsaQkFIRXLMrRDh2lwmOkyvZZtKkBHl1YTXUhXedzCVM5j_V-_qFj-mzqlx2reC4uHnmUVm7n5RblAoDI0o5wwlwfKDuK1M_OOoQY-Z7lxDI5ZOLOvWWy4vmsllQUsQgkm5sxCyZdJyNpZ26QWO-KuOcIbCsm0zV/w400-h266/2023-10-06T060942Z_867183703_RC2FL3AHA8TD_RTRMADP_3_SOUTHCHINASEA-PHILIPPINES-1200x800.jpg" width="400" /></a></div><br /><p class="MsoNormal">They haven’t rammed our vessels yet, but apparently feel no compunctions about
ramming the vessels of a nation perceived as weaker. This is classic bullying, just like what happens among school children. And only raw,
naked force stops it, just like
among school children. The difference is
that between nations, the force is
necessarily military action.<o:p></o:p></p><p class="MsoNormal">The worst offenders in recent years have been the
Chinese, the Russians, the Iranians,
and the North Koreans. You have
to understand who and what these players really are. All of these are brutal dictatorships
masquerading as faux democracies. Each
dictator is the center of his own fearless leader cult. Facts do not matter, and all opposition is crushed, creating the illusion of popular support. <o:p></o:p></p><p class="MsoNormal">These bullying dictators typically try to intimidate
neighbors into letting them do whatever they want, and if unopposed, the offensive behavior always escalates. If never opposed, the bully eventually loses his fear of
starting a war, and starts one. <o:p></o:p></p><p class="MsoNormal">Xi’s Chinese sail and fly too close to the ships and planes
of other nations in the ocean regions they want to annex. This has now escalated into ramming
Philippine ships. The Iranians under the
mullahs are seizing ships in the Straits of Hormuz and have now conducted a
major attack upon Israel with their proxy Hamas. Kim Jong Un’s North Koreans sank a South Korean ship a few
years ago, now they threaten dropping
atom bombs on the South, and on us and
Japan. It is all escalating. <o:p></o:p></p><p class="MsoNormal">Putin’s Russia first annexed part of Ukraine, then sent a proxy army to cause death and
destruction in eastern Ukraine, and still
being unopposed, Putin is losing his
fear of starting a war with NATO, and
has now invaded Ukraine, committing all
sorts of atrocities along the way. <o:p></o:p></p><p class="MsoNormal">If these ever-escalating bad behaviors sound familiar, they should.
Nazi Germany was a dictatorship built around Adolf Hitler as the center
of a fearless leader cult. His offensive
behaviors first included annexing the Rhineland, then the Sudentland, then all of Czechoslovakia and Austria. Never opposed, he lost his fear of war and invaded
Poland, starting World War 2 in
Europe. And we know all too well about
what happened as a result, and all the
atrocities he committed. <o:p></o:p></p><p class="MsoNormal">Xi’s China has also so far been unopposed with force. The next step for them is invading Taiwan, which could well start World War 3 in the
Pacific. Oh, you heard about that? Now you understand the seriousness of the pattern
here, because I just told you. <o:p></o:p></p><p class="MsoNormal">Diplomacy and sanctions are no opposition for these bullying
dictators. Only raw naked force gets
their attention. Without it, their escalating misbehavior does not
change. THAT is the lesson of history! Learn the lesson, or repeat the history! Simple as that!<o:p></o:p></p><p class="MsoNormal">The next time one of these dictator’s ships endangers one of
ours on the high seas, fire a shot
across its bow (the signal understood for centuries). If they don’t adjust, then sink them, no if’s,
and’s, or but’s. The next time one of their planes endangers
one of ours, fire a stream of tracers
across its path. If they don’t
adjust, shoot them down. It is that simple, and that brutal. There is no other way. <o:p></o:p></p><p class="MsoNormal">Global war is now upon you very soon, if you fail to learn the lesson of
history! <o:p></o:p></p><p class="MsoNormal">I have written about this problem before, at greater length than here. If you want more depth, go here on <a href="http://exrocketman.blogspot.com/">http://exrocketman.blogspot.com</a> and
see either (or both) the articles “What To Do About Aggressive Behavior At Sea
Or In the Air” dated 8 June 2023, and
“Bullying Demands Resistance”, dated 8
June 2019. There is a fast navigation
tool on the left side of that page.
Click on the year, then the
month, then the title if more than one article
was posted that month. <o:p></o:p></p><p class="MsoNormal">You will soon see that I have not changed my opinions about
what has been happening, or what is
required to stop it. The only thing
different is the increasing urgency as this escalates into war. <o:p></o:p></p><p>
</p><p class="MsoNormal">If you end up agreeing with my assessment, then contact your federal
representation, and tell them what you
want done. If they won’t do it, then elect somebody else. Simple as that. <o:p></o:p></p><p class="MsoNormal"><o:p></o:p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-87655597723574230492023-10-22T13:07:00.003-05:002023-11-30T10:25:43.425-06:00On the Israel-Hamas War<p><b><u>This was originally written 10-18-2023</u>:</b></p><p>To understand what is going on, you have to understand who and what the
players are. We know who the Israelis
are. But many do not understand who and
what Hamas (and Hezbollah) really are! We
need to include Hezbollah, because very
soon they might enter this war, too.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">These are terrorist organizations masquerading as religious
jihadist groups,<span style="mso-spacerun: yes;"> </span>that are really proxy
armies for Iran,<span style="mso-spacerun: yes;"> </span>and whose role is to
cause chaos and death wherever the Iranian government wants it.<span style="mso-spacerun: yes;"> </span>When I say “Iran”,<span style="mso-spacerun: yes;"> </span>I am talking about their government,<span style="mso-spacerun: yes;"> </span>a terrorist dictatorship masquerading as a
democracy.<span style="mso-spacerun: yes;"> </span>Not the Iranian people. <o:p></o:p></p>
<p class="MsoNormal">There are other similar terrorist organizations masquerading
as religious jihadist groups out there (Al Qaeda and ISIS come to mind),<span style="mso-spacerun: yes;"> </span>but it is Hamas and Hezbollah that belong to
the Iranian government who largely funds them,<span style="mso-spacerun: yes;">
</span>and certainly gives them training,<span style="mso-spacerun: yes;">
</span>weapons,<span style="mso-spacerun: yes;"> </span>and marching
orders.<span style="mso-spacerun: yes;"> </span>This has been going on for some
decades now,<span style="mso-spacerun: yes;"> </span>it is no secret. <o:p></o:p></p>
<p class="MsoNormal">I classify them as extreme terrorist groups,<span style="mso-spacerun: yes;"> </span>because that is exactly what they do.<span style="mso-spacerun: yes;"> </span>These groups kill anybody and everybody.<span style="mso-spacerun: yes;"> </span>In point of fact,<span style="mso-spacerun: yes;"> </span>they kill more Muslims than anybody
else,<span style="mso-spacerun: yes;"> </span>because those are the most
numerous potential victims in the countries where they typically operate.<span style="mso-spacerun: yes;"> </span>We’ve all seen these vicious crimes going on
for many years,<span style="mso-spacerun: yes;"> </span>and now Hamas terrorists
beheading little babies in this war.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">What else is this but simple murderous terrorism?<span style="mso-spacerun: yes;"> </span>It needs to be eliminated from the face of
the Earth,<span style="mso-spacerun: yes;"> </span>there is no question about
that! <o:p></o:p></p>
<p class="MsoNormal">The start of this current war has Iran’s fingerprints all
over it!<span style="mso-spacerun: yes;"> </span>Iran is a sworn enemy of
Israel,<span style="mso-spacerun: yes;"> </span>but also a sworn enemy of Saudi
Arabia (no secret there).<span style="mso-spacerun: yes;"> </span>Israel and
Saudi Arabia were getting close to a peace agreement that would have
effectively put severe restrictions on Iran’s freedom-of-action,<span style="mso-spacerun: yes;"> </span>so Iran had to start a war involving the
Israelis to put a stop to that.<span style="mso-spacerun: yes;"> </span>They did
that by commanding Hamas to attack in great force,<span style="mso-spacerun: yes;"> </span>which demands an invasion of the Gaza Strip
as the response.<o:p></o:p></p>
<p class="MsoNormal">Hamas has been attacking Israel from the Gaza Strip for many
years.<span style="mso-spacerun: yes;"> </span>This usually takes the form of an
unguided rocket barrage,<span style="mso-spacerun: yes;"> </span>or a few
terrorist troops crossing the border,<span style="mso-spacerun: yes;">
</span>but rarely both,<span style="mso-spacerun: yes;"> </span>and never before
in the quantities and coordination seen at the start of this war.<span style="mso-spacerun: yes;"> </span>That suggests the planning and supplies came
directly from Iran!<span style="mso-spacerun: yes;"> </span>Most of the weapons
do,<span style="mso-spacerun: yes;"> </span>we’ve already seen some stockpiles
of them on TV.<o:p></o:p></p>
<p class="MsoNormal">As for the explosion at the hospital in southern Gaza,<span style="mso-spacerun: yes;"> </span>consider that logically there are 4
possibilities:<span style="mso-spacerun: yes;"> </span>(1) Israel did this
deliberately,<span style="mso-spacerun: yes;"> </span>(2) Israel did this
accidentally,<span style="mso-spacerun: yes;"> </span>(3) Hamas did this
deliberately,<span style="mso-spacerun: yes;"> </span>and (4) Hamas did this
accidentally.<span style="mso-spacerun: yes;"> </span>The point of such an
attack,<span style="mso-spacerun: yes;"> </span>if deliberate,<span style="mso-spacerun: yes;"> </span>would be to stir up anti-Israeli sentiment
among not just the Arabs,<span style="mso-spacerun: yes;"> </span>but around the
world.<span style="mso-spacerun: yes;"> </span>Which we have seen!<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">We have not seen any evidence as of yet,<span style="mso-spacerun: yes;"> </span>regarding which of the 4 possibilities might
really be true.<span style="mso-spacerun: yes;"> </span>We have seen some
suggestions that this was a misfired Hamas rocket (possibility 4). But no
evidence has shown up yet in public.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">Instead consider motivations.<span style="mso-spacerun: yes;"> </span>Why would Israel want such a thing to
happen?<span style="mso-spacerun: yes;"> </span>It could only stir up resentment
against them.<span style="mso-spacerun: yes;"> </span>Hamas,<span style="mso-spacerun: yes;"> </span>on the other hand,<span style="mso-spacerun: yes;"> </span>would like that resentment against the
Israelis.<span style="mso-spacerun: yes;"> </span>That suggests possibilities 1
and 2 are rather unlikely,<span style="mso-spacerun: yes;"> </span>while
possibilities 3 and 4 are rather likely. <o:p></o:p></p>
<p class="MsoNormal">So the odds favor that Hamas did this,<span style="mso-spacerun: yes;"> </span>deliberate or not.<span style="mso-spacerun: yes;"> </span>I cannot rule out deliberate,<span style="mso-spacerun: yes;"> </span>because Hamas is the elected government (such
as it is) of Gaza.<span style="mso-spacerun: yes;"> </span>And yet Hamas has a repeatedly-demonstrated
proclivity to use as human shields the very people it supposedly governs!<span style="mso-spacerun: yes;"> </span><o:p></o:p></p><p class="MsoNormal"></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgeb4LAfAd3yU_pNWbkqFYWMBgZTLQcWEBrck-8Z3fPK-EhmVwFt4HopEhf0yFhFLWYKygrb_fojYgg4gpAtVrMChPnCSLg8SyEG6tZS3ARiuamVNwB3Zl4nNiHjHm9x7aV79vMAM-cH7NbXUzUz-CWKSGeeMwSjoEYgOMJKrfOz0tD1U66t_O17tqXVWLy/s1200/hamas_infographic-1200x1190.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1190" data-original-width="1200" height="396" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgeb4LAfAd3yU_pNWbkqFYWMBgZTLQcWEBrck-8Z3fPK-EhmVwFt4HopEhf0yFhFLWYKygrb_fojYgg4gpAtVrMChPnCSLg8SyEG6tZS3ARiuamVNwB3Zl4nNiHjHm9x7aV79vMAM-cH7NbXUzUz-CWKSGeeMwSjoEYgOMJKrfOz0tD1U66t_O17tqXVWLy/w400-h396/hamas_infographic-1200x1190.png" width="400" /></a></div><p></p><p class="MsoNormal">Israel did not expect an attack on this scale. That much is clear, because it was absolutely not prepared
ahead-of-time to invade. In the
past, an air raid or two was sufficient
response. But not this time! <o:p></o:p></p><p class="MsoNormal">I suggest that they invade,
kill all the Hamas fighters they can find, and corral the civilians. The women-with-children and old men can
pretty much be presumed innocent civilians.
Younger men in the company of women-with-children are likely
civilians, too. But young single men, and single-women-without-children (a rarity
in typical Arab cultures), need close
scrutiny! <o:p></o:p></p><p class="MsoNormal">Why? Because you do
not want to kill innocent civilians, but
you do not want to take any Hamas prisoners!
Why waste the resources to keep alive that which you so sincerely want
dead? And if Hezbollah comes across the
border from Lebanon, don’t take any of
them prisoners, either! This scourge needs to be wiped clean from the
face of the Earth. <o:p></o:p></p><p class="MsoNormal"><b><u>Update 10-22-2023</u>:<o:p></o:p></b></p><p class="MsoNormal">4 days after this was originally written, we have still seen no evidence in public
about who really blew up that hospital in southern Gaza. The motivations still point to Hamas doing it
either deliberately or accidentally, and
some reports finger a subsidiary Hamas terrorist group as the perpetrators. <o:p></o:p></p><p class="MsoNormal">On the other hand,
terrorists are now firing rockets at Israel from the West Bank, and Syria,
if reports are accurate. Everyone
fears a widening war. It definitely
appears that Hamas and its sympathizers are the ones widening it. <o:p></o:p></p><p class="MsoNormal">The recommended solution is still the same: wipe these evil people from the face of the
Earth. All of them. But,
while doing it, try not to kill
the innocent civilians!<o:p></o:p></p><p class="MsoNormal">
<span style="mso-spacerun: yes;"></span></p><p class="MsoNormal"><span style="mso-spacerun: yes;"><b><u>Update 11-2-2023</u>:<o:p></o:p></b></span></p><span style="mso-spacerun: yes;">
<p class="MsoNormal"><a name="_Hlk149815662">Israeli forces are now in the Gaza
strip, <span style="mso-spacerun: yes;"> </span>surrounding Gaza city, <span style="mso-spacerun: yes;"> </span>and hitting what they claim are Hamas targets
with bombing,<span style="mso-spacerun: yes;"> </span>while causing casualties
among the civilians in and around these targets.<span style="mso-spacerun: yes;"> </span>Meanwhile Hamas is still firing rockets into
Israel,<span style="mso-spacerun: yes;"> </span>and so is Hezbollah in
Lebanon.<span style="mso-spacerun: yes;"> </span>Plus,<span style="mso-spacerun: yes;"> </span>other Iranian terrorist proxies,<span style="mso-spacerun: yes;"> </span>are firing directly at American forces in
Iraq and Syria,<span style="mso-spacerun: yes;"> </span>and killing some of them.<span style="mso-spacerun: yes;"> </span><b><i>The war <u>has already widened</u>,<span style="mso-spacerun: yes;"> </span>whether you like it,<span style="mso-spacerun: yes;"> </span>or not!</i></b><span style="mso-spacerun: yes;"> </span>There are two fundamental points to make
about this,<span style="mso-spacerun: yes;"> </span>which so far are not being
addressed properly in the press reporting. <o:p></o:p></a></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;"><b>Point 1:</b><span style="mso-spacerun: yes;"> </span><b>the ultimate culprit here is the terrorist
“government” of Iran.</b><span style="mso-spacerun: yes;"> </span>All these violent
groups are their proxy terrorist armies.<span style="mso-spacerun: yes;">
</span>The Iranian government funds them,<span style="mso-spacerun: yes;">
</span>supplies them,<span style="mso-spacerun: yes;"> </span>and gives them
strategic direction.<span style="mso-spacerun: yes;"> </span>So far,<span style="mso-spacerun: yes;"> </span>the government of Iran has not been held
accountable for any of this destruction,<span style="mso-spacerun: yes;">
</span>despite their being fundamentally at fault for all of it!<span style="mso-spacerun: yes;"> </span>That is a very serious foreign policy mistake
being made by all the western governments,<span style="mso-spacerun: yes;">
</span>collectively.<span style="mso-spacerun: yes;"> </span><b><i>And this evil
crap will not stop until the terrorist government of Iran is deposed! </i></b><span style="mso-spacerun: yes;"> </span><o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;">That government is
terrorist dictatorship that masquerades as a religious group to justify itself
to the people it rules,<span style="mso-spacerun: yes;"> </span>and it operates
a sham democracy where the ruling “mullahs” can override anything any elected
body decides.<span style="mso-spacerun: yes;"> </span>They ruthlessly oppress
and kill any of their own citizens who dare protest or defy anything. This has
been going on since they took over in 1979 after the Shah fled.<span style="mso-spacerun: yes;"> </span><o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;"><u>You’d think the
western governments would have done something effective about this before
now,<span style="mso-spacerun: yes;"> </span>but they have not</u>.<o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;"><b>Point 2:<span style="mso-spacerun: yes;"> </span>the calls for a cease fire are
well-meaning,<span style="mso-spacerun: yes;"> </span>but are strategically very
incorrect.</b><span style="mso-spacerun: yes;"> </span>The reporting on this is
based upon the massive death and destruction of civilians in Gaza that is right
in front of everyone,<span style="mso-spacerun: yes;"> </span>and not at all on <u>what
underlies that tragedy</u>,<span style="mso-spacerun: yes;"> </span>and which overwhelms
that tragedy in importance.<span style="mso-spacerun: yes;"> </span><o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;">That mistake on the
part of most of the reporters is understandable,<span style="mso-spacerun: yes;"> </span>since reporting on what is right in front of
them, <span style="mso-spacerun: yes;"> </span>is usually exactly what they are
supposed to do.<span style="mso-spacerun: yes;"> </span>But,<span style="mso-spacerun: yes;"> </span>focusing on the appearance is not what must
be done by those who actually have to fight the war!<span style="mso-spacerun: yes;"> </span>The few reporters I have seen who actually
interview a real war-fighter about this issue,<span style="mso-spacerun: yes;">
</span>do not seem to understand,<span style="mso-spacerun: yes;"> </span>or
even listen to,<span style="mso-spacerun: yes;"> </span>the answers given by
these war-fighters,<span style="mso-spacerun: yes;"> </span>since these answers
disagree with the appearances. <span style="mso-spacerun: yes;"> </span>And that
is a very serious mistake being made by the news media.<span style="mso-spacerun: yes;"> </span>Dig deeper,<span style="mso-spacerun: yes;">
</span>please!<o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;">Unlike previous Israeli
responses to Hamas attacks,<span style="mso-spacerun: yes;"> </span>this time
the Israelis have made their top objective the ultimate destruction of
Hamas,<span style="mso-spacerun: yes;"> </span>despite the human shield behind
which Hamas is hiding.<span style="mso-spacerun: yes;"> </span>Why should it
surprise anyone that there are civilian casualties,<span style="mso-spacerun: yes;"> </span>when to win,<span style="mso-spacerun: yes;">
</span>the Israelis must go <u>right through that human shield</u> to get at
Hamas?<span style="mso-spacerun: yes;"> </span>If they do not do this,<span style="mso-spacerun: yes;"> </span>they lose,<span style="mso-spacerun: yes;">
</span>and so do we all,<span style="mso-spacerun: yes;"> </span>since the world
must be rid of these terrorist groups!<span style="mso-spacerun: yes;"> </span><b><i>The
real war crime here is not the Israelis going through that human shield,<span style="mso-spacerun: yes;"> </span><u>it is Hamas using a human shield in the
first place</u>!</i></b><span style="mso-spacerun: yes;"> </span><o:p></o:p></span></p>
<p class="MsoNormal"><span style="mso-bookmark: _Hlk149815662;"><u>And I do not see
that crucial fact being adequately reported</u>.<o:p></o:p></span></p><b><u>Update 11-6-2023</u>:</b> For those who cannot see past the civilian casualties in Gaza, consider this. Israel telegraphed the "punch" that was coming, to give the civilians who constituted the human shield time to get out of the northern Gaza strip. Very few did, most did not. Some might have since changed their minds, worthy of a short pause to depart, but only very short! The Hamas terrorists simply cannot be allowed time to regroup. </span><div><b><u><br /></u></b></div><div><b><u>Update 11-30-2023</u>:</b></div><div>
<p class="MsoNormal">Please see the column by Jonah Goldberg and titled “China
deflects world’s attention to Israel”,
which was published this date in the Waco Tribune-Herald. He hits the nail right on the head regarding
the strategy China has been using to deflect away criticisms for the many evils
it does. <o:p></o:p></p>
<p class="MsoNormal">The misplaced support so many very young people show for the
terrorist “government” in Gaza, called
Hamas, traces directly to their use of
the Tik Tok social media platform, where
anti-Jewish disinformation (among other things) gets spread by the Chinese
government. China is not the only one
doing this, but right now, they are among the worst.<o:p></o:p></p>
<p class="MsoNormal">Thank you, Jonah Goldberg,
for pointing this danger out in public.<o:p></o:p></p><p></p><div><span style="mso-spacerun: yes;"><br /></span></div></div>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com4tag:blogger.com,1999:blog-2675974463524895416.post-30199811085142939812023-10-20T16:36:00.003-05:002023-10-20T16:36:41.832-05:00Scariest Halloween Costume<p> </p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgjPdB28AH4GXrPUEI4JJvJqF2u-WcxarFe7n3KSdljdRa9NKCQ2Fz3DTHMMfzeR-j2SnVwI6jo0fC3cJq549zrD18B0qEvNJYZE9Ro0kxRnDXZ_YWIhITbYL5aaY7yMgFY7bqSnL8Tjsd1T8swB0sThzsLqKArdrrrQ8z1SS93tbbz9WsEMS0FHhnAmHuI/s643/halloween%20T%20shirt.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="522" data-original-width="643" height="325" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgjPdB28AH4GXrPUEI4JJvJqF2u-WcxarFe7n3KSdljdRa9NKCQ2Fz3DTHMMfzeR-j2SnVwI6jo0fC3cJq549zrD18B0qEvNJYZE9Ro0kxRnDXZ_YWIhITbYL5aaY7yMgFY7bqSnL8Tjsd1T8swB0sThzsLqKArdrrrQ8z1SS93tbbz9WsEMS0FHhnAmHuI/w400-h325/halloween%20T%20shirt.png" width="400" /></a></div><br /><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-20196744916664544062023-10-18T19:03:00.006-05:002023-10-25T16:04:24.446-05:00Jim Jordan, Speaker of the House? NO!!!<p>I’m sorry, but the
record established by Jim Jordan as a representative from Ohio is much more an
agent of chaos, than anything one might
construe as doing the people’s business in Congress. He is an ardent Trump supporter, a 2020 election denier, and there is at least some reason to believe
he should also be under indictment, regarding the Jan 6, 2021,
insurrection at the Capitol.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal"><b><i>This is absolutely NOT the man you want running the
agenda for the House of Representatives,<span style="mso-spacerun: yes;">
</span>who is also third in line to Presidency!<o:p></o:p></i></b></p>
<p class="MsoNormal">What you are seeing here is the Trump cult loud minority
trying to find a way to rig things so that they can pull the levers of power in
Washington,<span style="mso-spacerun: yes;"> </span>despite not having the votes
to get their way properly.<span style="mso-spacerun: yes;"> </span>Failing
that,<span style="mso-spacerun: yes;"> </span>I predict (<b><i>warn!!!</i></b>)
that they will continue to cause paralysis in the House,<span style="mso-spacerun: yes;"> </span>causing the government to be
dysfunctional,<span style="mso-spacerun: yes;"> </span>and drive it into default
on its debt.<span style="mso-spacerun: yes;"> </span><u>This is deliberate
sabotage</u>!<o:p></o:p></p>
<p class="MsoNormal">The Trump cultists are hoping there will be an outcry from
the people to “do something” if this chaos persists long enough and causes
enough destruction.<span style="mso-spacerun: yes;"> </span>Their hope is to
topple our democracy and install a dictatorship in its place,<span style="mso-spacerun: yes;"> </span>something espoused in public by former Trump
advisor retired General Flynn (<u>many have forgotten that,<span style="mso-spacerun: yes;"> </span>but I did not</u>!),<span style="mso-spacerun: yes;"> </span>with Trump (or a younger Trump wannabee like Desantis) as
dictator. <o:p></o:p></p>
<p class="MsoNormal">This is <b>absolutely NOT</b> the Republican party that I
once knew and respected!<span style="mso-spacerun: yes;"> </span>There are now
no longer enough traditional Republicans to resist the Trump cultists.<span style="mso-spacerun: yes;"> </span>This Trump cult group is now a direct,<span style="mso-spacerun: yes;"> </span>clear,<span style="mso-spacerun: yes;">
</span>and present danger to our republic.<span style="mso-spacerun: yes;">
</span>Personally,<span style="mso-spacerun: yes;"> </span>I see little
difference between them,<span style="mso-spacerun: yes;"> </span>and what
happened in 1933 Germany with the Nazi takeover from within. The Nazis were a Hitler cult.<o:p></o:p></p>
<p class="MsoNormal"><b><i>Do not be fooled by what they say.<span style="mso-spacerun: yes;"> </span>Look instead at what they actually do!<span style="mso-spacerun: yes;"> </span>Which speaks volumes!<span style="mso-spacerun: yes;"> </span><o:p></o:p></i></b></p>
<p class="MsoNormal">My sincere recommendation:<span style="mso-spacerun: yes;">
</span>vote these people out,<span style="mso-spacerun: yes;"> </span>as fast
and as decisively as you can!<span style="mso-spacerun: yes;"> </span>It does
not matter who the opponent is, <span style="mso-spacerun: yes;"> </span>or what
his party affiliation is.<span style="mso-spacerun: yes;"> </span>You cannot do
any worse than what you have right now!<span style="mso-spacerun: yes;">
</span>You might do better!<o:p></o:p></p>
<p class="MsoNormal">You CANNOT say that I did not warn you!<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhd7D6IXSdS3dApSaJkNLTvV3y1aW9huJcPXxJJHeC5fXbvxLSGGzpYnxbfRjNHTH0jzsRTEx-rab2ViSb924VGqpCKZfbTCHQ76OQaJnqG027KPq6bpLQsqUXOYRKjwxiAp3u-2cfRvWjKusJB47Wni7JIlNU7A4EAXjFN_t-_D-w-EuiJSlXUyLQ44Nl7/s273/Jordan.jpg" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="158" data-original-width="273" height="232" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhd7D6IXSdS3dApSaJkNLTvV3y1aW9huJcPXxJJHeC5fXbvxLSGGzpYnxbfRjNHTH0jzsRTEx-rab2ViSb924VGqpCKZfbTCHQ76OQaJnqG027KPq6bpLQsqUXOYRKjwxiAp3u-2cfRvWjKusJB47Wni7JIlNU7A4EAXjFN_t-_D-w-EuiJSlXUyLQ44Nl7/w400-h232/Jordan.jpg" width="400" /></a></div><br /><p class="MsoNormal"><b><u>Update 10-20-2023</u>: </b>Jordan lost a 3rd vote (with fewer votes each time) and has been deleted as a GOP candidate for Speaker, according to news reports. Whew, that's a relief! Now we have to see who they might nominate instead. The risk would be another Trump-cult agent of chaos. </p><p class="MsoNormal"></p><p class="MsoNormal"><b><u>Update 10-23-2023</u>:</b><span style="mso-spacerun: yes;"> </span>The House GOP is now considering 9 candidates
from which to select a nominee to be speaker.<span style="mso-spacerun: yes;">
</span>I checked the news stories to find out who they are and a little bit
about them.<span style="mso-spacerun: yes;"> </span>7 of the 9 voted to
decertify the 2020 election results <u>after</u> the Jan 6 2021 insurrection
took place!<span style="mso-spacerun: yes;"> </span>Of the other 2,<span style="mso-spacerun: yes;"> </span>one supported the lawsuit in Texas to
overturn the 2020 election results (Emmer),<span style="mso-spacerun: yes;"> </span>and
the other was a Jim Jordan supporter (Scott), <span style="mso-spacerun: yes;"> </span>right up until he dropped out.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">As far as I am concerned,<span style="mso-spacerun: yes;">
</span>the ties all 9 candidates have to the Trump-cult attempt to overturn the
election and stop the peaceful transfer of power,<span style="mso-spacerun: yes;"> </span><span style="mso-spacerun: yes;"> </span><u>totally
disqualifies each and every one of them</u> as a credible candidate to be
Speaker of the House!<span style="mso-spacerun: yes;"> </span>They only have
credibility with the Trump cultists.<span style="mso-spacerun: yes;"> </span>I
do not see in any of them, <span style="mso-spacerun: yes;"> </span>a candidate
that even might possibly prioritize actually doing the people’s business, <span style="mso-spacerun: yes;"> </span>above their extremist political agenda. <o:p></o:p></p>
<p class="MsoNormal"><b><i>My conclusion from all that is this:</i></b><span style="mso-spacerun: yes;"> </span>any non-Trump-cult Republicans that still might
remain in the House need to get with the House Democrats and do the suggested
bipartisan deal.<span style="mso-spacerun: yes;"> </span>I think we will soon
see if any such still exist in the House.<span style="mso-spacerun: yes;">
</span>If not,<span style="mso-spacerun: yes;"> </span>get ready for
long-lasting paralysis of Congress, <span style="mso-spacerun: yes;"> </span>and
a debt payment default that hurts us all,<span style="mso-spacerun: yes;">
</span><b><i>plus a total calamity in our foreign policy during a time of
multiple crises plus a war that seriously threatens to widen and engulf us all</i></b>.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">That last <b><u>aids and comforts our enemies</u></b>,<span style="mso-spacerun: yes;"> </span>and <b><u>that</u></b> is in the
Constitutional definition of treason!<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal"><u>Merrick Garland and the DOJ</u>:<span style="mso-spacerun: yes;"> </span>are you listening?<span style="mso-spacerun: yes;"> </span>Did you hear what I just said?<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">You need to do something about Trump-cult treason,<span style="mso-spacerun: yes;"> </span>and you need to do it quickly!!! <span style="mso-spacerun: yes;"> </span>You will have millions of eyewitnesses to these
acts of treason,<span style="mso-spacerun: yes;"> </span>witnesses who will have
seen this on live TV<b>. <u>One or more eyewitnesses
to the act</u></b> is a Constitutional requirement,<span style="mso-spacerun: yes;"> </span>too. <span style="mso-spacerun: yes;"> </span>And it will be satisfied. <o:p></o:p></p><b><u>Update 10-25-2023</u>:</b> They selected Emmer as their candidate, but he dropped out only hours afterward, when both Trump and the Trump-cult extremists denounced him (as not being enough of a Trump supporter). Now they have selected Mike Johnson of Louisiana, one of the 7 who voted to decertify the election, and a well-known ardent Trump supporter and Jordan supporter. We will see where that goes! <div><br /></div><div>The choice seems to be to either give control of the House to the Trump-cult extremists (leading directly to the end of our democracy), or simply continue the ongoing chaos and paralysis (leading perhaps to the end of our democracy). Which is EXACTLY what I warned you readers about above! <p></p></div><div><b><u>Late-breaking same day</u>: </b> the non-Trump-cult Republicans gave up and let the House elect Mike Johnson of Louisiana as its Speaker. That gives total control of 1 house of Congress to a bunch of radical extremists that are bound and determined to destroy our democracy and replace it with a dictatorship. Expect nothing useful from Congress, and thus our entire government, from this point! </div><div><br /></div><div>Expect to see some non-Trump-cult Republicans assassinated between now and the 2024 election. It's called a "purge". We've seen this before, with the Fascists, the Nazis, and the Bolsheviks, as well as the Maoists. </div><div><br /></div><div>Expect also that they will attempt to install Trump as dictator, via the 2024 election, whether he is a convicted felon or not, and whether he is behind bars or not! </div><div><br /></div><div>Better arm yourselves! It is only a matter of time before these extremists come for you, too! </div>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-73943597868297241862023-10-06T09:17:00.002-05:002023-10-06T09:17:47.136-05:00Two Funnies<p>Here are two funnies that my wife found on Facebook, presented essentially without comment. Enjoy! (Youngsters might need to be reminded that the old man in the second picture is Dick Van Dyke, who is 97 years old now.)</p><p><br /></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhrXDSAfflJBP0ohlOTAIbXHaYaLCZDKuL2BRoWpWw2dMeXHx1iKVtxsnxF7J9Z5FogI_Ectz-Fs-NXNRu-lMSTDzKXunKUV4rEGGgmeHBBo-NlwBRvzFBlMGW2aRMqh7rAkzLIES0eCzTt39FkjYtWXVHd8KiuqI7pRD_sZCa64hE-Ond7JxuU_myced6e/s1024/warming.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="998" data-original-width="1024" height="390" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhrXDSAfflJBP0ohlOTAIbXHaYaLCZDKuL2BRoWpWw2dMeXHx1iKVtxsnxF7J9Z5FogI_Ectz-Fs-NXNRu-lMSTDzKXunKUV4rEGGgmeHBBo-NlwBRvzFBlMGW2aRMqh7rAkzLIES0eCzTt39FkjYtWXVHd8KiuqI7pRD_sZCa64hE-Ond7JxuU_myced6e/w400-h390/warming.png" width="400" /></a></div><br /><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1RuMu7qPG4T4x4IKJjMUKZjDDuGh4oBuDK9qEXQPcP_jWAcy-2E_yVLQWsiLr2E1OrqNWJtnrKJ6ul70i_g8bvMbn5KQak-AjamwJeld0_wtYal8-bY68XJvy2o1bEVMo_E24aOVsofFFeUZP8_OSk3s6pkdGUnH1uefpG1ytA5XPbJJNpzBz2aqxUM1H/s1188/names.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1188" data-original-width="1178" height="400" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi1RuMu7qPG4T4x4IKJjMUKZjDDuGh4oBuDK9qEXQPcP_jWAcy-2E_yVLQWsiLr2E1OrqNWJtnrKJ6ul70i_g8bvMbn5KQak-AjamwJeld0_wtYal8-bY68XJvy2o1bEVMo_E24aOVsofFFeUZP8_OSk3s6pkdGUnH1uefpG1ytA5XPbJJNpzBz2aqxUM1H/w396-h400/names.png" width="396" /></a></div><br /><p><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-83212189060848384792023-10-01T11:26:00.001-05:002023-10-01T13:41:24.622-05:00Basic Thermal Results for High Speeds<p class="MsoNormal">This article is a direct follow-on <u>to the updates posted
to</u> “Purported SR-72 Propulsion”,
posted 1 September 2023. As I
have said there, and multiple places and
times elsewhere, <b><i>if you do not
have a thermal management design concept,
you do not have a feasible hypersonic flight concept!</i></b> This article attempts to put some bounds on
that problem. <o:p></o:p></p>
<p class="MsoNormal"><b>Lateral Skin Study<o:p></o:p></b></p>
<p class="MsoNormal">The following is a simplified equilibrium skin panel surface
temperature estimate for lateral-facing skin panels. These could be on aerosurfaces (wings and
fins), or on the sides of a fuselage
body. I did not consider any conduction
inward or to adjacent panels. I did not
consider any active cooling. There is
convection to the panel, and thermal
re-radiation from it. It soaks out hot
enough to balance the two. <o:p></o:p></p>
<p class="MsoNormal">I did this for Mach numbers from subsonic to Mach 7, using standard compressible flow methods and
the high-speed heat transfer models that are based upon it. I used free-stream conditions as the good
approximation that they really are, for
local edge-of-boundary layer conditions.
I did not analyze past Mach 7,
because the fundamental assumptions underlying compressible flow
analysis methods are breaking down, due
to ionization into something that is no longer air as we know it. <o:p></o:p></p>
<p class="MsoNormal"><o:p></o:p></p><p>I show temperature curves in <b>Figure 1</b> for air total
temperature, boundary layer recovery temperature
(the driver for heat transfer to the panel),
and equilibrium panel soak temperatures for low and high thermal
emissivity. The service temperature
limits for a variety of materials are also shown. <b>Figure 2</b> shows the film coefficient
trends vs Mach at 40 kft, for low and
high emissivity. Beyond about Mach 3 or
4, these are pretty constant. Data in the same formats for 85 kft are in <b>Figure
3 and 4</b>, and for 130 kft <b>Figures
5 and 6</b>. </p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgFnuiIWWx9wHLDAwHrRSdocg8Hy8Hls4sAWniXM9OcLFiuh1-M2fuxBIYX1yaFgs68gIacgXNSBqVtdCmdnw_AaVCZglTkABBRSs7CUtbptl_eIYQ8_H5u7tTXAA6WvDaIdYvJw1h-iOXZrYjMS5Cg9KTqXuYM51TeR2VuBKJebbzYYvOH9B57SVFeK6rA/s1044/F1%2040%20kft%20demo.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="589" data-original-width="1044" height="226" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgFnuiIWWx9wHLDAwHrRSdocg8Hy8Hls4sAWniXM9OcLFiuh1-M2fuxBIYX1yaFgs68gIacgXNSBqVtdCmdnw_AaVCZglTkABBRSs7CUtbptl_eIYQ8_H5u7tTXAA6WvDaIdYvJw1h-iOXZrYjMS5Cg9KTqXuYM51TeR2VuBKJebbzYYvOH9B57SVFeK6rA/w400-h226/F1%2040%20kft%20demo.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 1 – Skin Panel Soak-Out vs Mach at 40 kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgicRPwN-Q74IaCKJBZkHffKweBiLm9rkNjsQdV3_ROFcnF4Y34zR-hQv-OPzBDVsvU50WxBXYpLSJHTbjiOZntpExyHGirDRo3TAD7WZNRBT4AwJPNlFuxufWXWC9v3_OenQbtVWKbD4lWSlQp_aaL46UFUbn3iUbZdSs1QyTaa-__ti-z0_5GG3dYlgW3/s1094/F2%20film%2040%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="573" data-original-width="1094" height="210" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgicRPwN-Q74IaCKJBZkHffKweBiLm9rkNjsQdV3_ROFcnF4Y34zR-hQv-OPzBDVsvU50WxBXYpLSJHTbjiOZntpExyHGirDRo3TAD7WZNRBT4AwJPNlFuxufWXWC9v3_OenQbtVWKbD4lWSlQp_aaL46UFUbn3iUbZdSs1QyTaa-__ti-z0_5GG3dYlgW3/w400-h210/F2%20film%2040%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 2 – Film Coefficients vs Mach at 40 kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjvwal2LOPqXn7fyO2Mrt6XxOgg0ZQfG16C72r6d7kUUNbJHuflEiYFEqniKuWvQIt8zDuIrR62M0WV9ORs_i8OHid9EWJxFAskJkdllmSxIyfOxniEhA1NPWVITPIzcmXUn6dI5T36kUUTYJXqGug_K0y45GWfUP_f9Uu-bZCOnRzuPfRioYRDmBjEyKCn/s1014/F3%2085%20kft%20student.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="616" data-original-width="1014" height="243" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjvwal2LOPqXn7fyO2Mrt6XxOgg0ZQfG16C72r6d7kUUNbJHuflEiYFEqniKuWvQIt8zDuIrR62M0WV9ORs_i8OHid9EWJxFAskJkdllmSxIyfOxniEhA1NPWVITPIzcmXUn6dI5T36kUUTYJXqGug_K0y45GWfUP_f9Uu-bZCOnRzuPfRioYRDmBjEyKCn/w400-h243/F3%2085%20kft%20student.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 3 – Skin Panel Soak-Out vs Mach at 85 kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi09e3TBtMTAm-dxytK_2Bc6qA3mNX7UPiwdOlNfxyRB_MUgLmqk4oM2EhVz6tjlU3KEp6XK1qHCh1N4fFm3tkCQibdLDURQ5PeIhdpdezTdbseY25XYrH5XPS8a-OCO6DIiIkfGUd-XMZC2mg-5Nyqtqfpnw5bDS00b1uIwWq9PwD9HpR0Kui2Mbe0PhcS/s1083/F4%20film%2085kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="623" data-original-width="1083" height="230" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEi09e3TBtMTAm-dxytK_2Bc6qA3mNX7UPiwdOlNfxyRB_MUgLmqk4oM2EhVz6tjlU3KEp6XK1qHCh1N4fFm3tkCQibdLDURQ5PeIhdpdezTdbseY25XYrH5XPS8a-OCO6DIiIkfGUd-XMZC2mg-5Nyqtqfpnw5bDS00b1uIwWq9PwD9HpR0Kui2Mbe0PhcS/w400-h230/F4%20film%2085kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 4 – Film Coefficients vs Mach at 85 kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjleAeM1HUwoPVMjYBfTszR0o_2vM9AXfvGFbR7H6918I_mAyQEqVXPfeHGODsumuYgvpeFbThvkbog2Udu_1wEjhxqeUffrmhtT6crzhrkmeP1bHMsAho_tPoto3tj1CEHM5jb-qrBzb_lEuhtwHdyZGIRmfDIrwEcnkPEdcWBlZ-OihWsiR2ET5cw4JCW/s1015/F5%20130%20kft%20student.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="606" data-original-width="1015" height="239" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjleAeM1HUwoPVMjYBfTszR0o_2vM9AXfvGFbR7H6918I_mAyQEqVXPfeHGODsumuYgvpeFbThvkbog2Udu_1wEjhxqeUffrmhtT6crzhrkmeP1bHMsAho_tPoto3tj1CEHM5jb-qrBzb_lEuhtwHdyZGIRmfDIrwEcnkPEdcWBlZ-OihWsiR2ET5cw4JCW/w400-h239/F5%20130%20kft%20student.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 5 – Skin Panel Soak-Out vs Mach at 130 kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhZwtOTKEzAjddoro5Xxx-EqhRS9LQcUzQX3AR8eYWqa-K8J6xs-hstOIKm7rD2KoiO1AGJuNVrHaGdTe5p1EvJr3c2J4W6bT1xY5HbQyFmk2S51lVGojBQfhfohINtilfSWkOv_ujTVnsPvbWAP8LP90DJW7dL71lEgXf4UF8bIVHQHh2DTHA771TGxMgJ/s1085/F6%20film%20130%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="628" data-original-width="1085" height="231" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhZwtOTKEzAjddoro5Xxx-EqhRS9LQcUzQX3AR8eYWqa-K8J6xs-hstOIKm7rD2KoiO1AGJuNVrHaGdTe5p1EvJr3c2J4W6bT1xY5HbQyFmk2S51lVGojBQfhfohINtilfSWkOv_ujTVnsPvbWAP8LP90DJW7dL71lEgXf4UF8bIVHQHh2DTHA771TGxMgJ/w400-h231/F6%20film%20130%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 6 – Film Coefficients vs Mach at 130 kft<o:p></o:p></p><p class="MsoNormal"><b>Skin Study Correlation:<o:p></o:p></b></p><p class="MsoNormal">Recovery temperatures do not change so drastically with
altitude, unlike film coefficients. See <b>Figure 7</b>.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgCHXKmq483HrqfwxEQTqeR27pxycrymGKyXV-r09iEiY8RuVo1rQNucIyf1IWg6tb62thUUySmTsArZNILAUj-Z4dMDhtJA1n8Jr43keZoctjeHWAOFW3xMxg0kl91Gg9i7ziI1EvpLlAU-pSsSU7jf0OCCgHueZxYGdtUklfzZ7I406pcbQDTyOD3-Vq9/s996/F7%20conclusion%201.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgCHXKmq483HrqfwxEQTqeR27pxycrymGKyXV-r09iEiY8RuVo1rQNucIyf1IWg6tb62thUUySmTsArZNILAUj-Z4dMDhtJA1n8Jr43keZoctjeHWAOFW3xMxg0kl91Gg9i7ziI1EvpLlAU-pSsSU7jf0OCCgHueZxYGdtUklfzZ7I406pcbQDTyOD3-Vq9/w400-h236/F7%20conclusion%201.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 7 – Replots of Film Coefficient and Soakout vs
Altitude at Mach 5<o:p></o:p></p><p class="MsoNormal">As the figure shows,
the result is a drastic change in soakout temperatures, driven by drastically lower film coefficients
at extreme altitudes. The recovery
temperatures all fall between 3800 and 4500 F at Mach 7, as shown in <b>Figures 1, 3, and 5</b> above. This suggests that a single analysis could
establish a representative film coefficient value insensitive to changes in
speed, at Mach 4+ and some altitude, which could be quickly scaled to other
altitudes. Calculating recovery
temperatures at each flight condition is a far easier thing to do. The correlation supporting that shortcut is
given in <b>Figure 8</b>. Doing it that
way is only a ballpark estimate that supports better, more detailed analyses later. But it is useful. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjWBCqchFdi2UKKfhNj_2rNu-MjxzPtGhBgVs1da0bUp6aIObccxDrlpKBxKBhBEV14WOqGHnuSuZMEWVqg5V8833UY3ShUphV_J9wl0XkFGpDDaf5dO-TM5nR2WflG252sbhuZB8SIufdgwnLvgJi9FFjbNukz6mk_72qHY_nQ3UkRFovkf5-ATWpNjkxO/s996/F8%20conclusion%202.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjWBCqchFdi2UKKfhNj_2rNu-MjxzPtGhBgVs1da0bUp6aIObccxDrlpKBxKBhBEV14WOqGHnuSuZMEWVqg5V8833UY3ShUphV_J9wl0XkFGpDDaf5dO-TM5nR2WflG252sbhuZB8SIufdgwnLvgJi9FFjbNukz6mk_72qHY_nQ3UkRFovkf5-ATWpNjkxO/w400-h236/F8%20conclusion%202.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 8 – Correlating High-Speed Film Coefficient vs
Altitude<o:p></o:p></p><p class="MsoNormal"><b>Leading Edge Stagnation Study<o:p></o:p></b></p><p class="MsoNormal">There is a compressible flow-based heat transfer correlation
for stagnation zone heating. It exists
in two forms, determined by a coefficient
on the Nusselt number expression: C =
1.28 for nose tips, and C = 0.95 for
aerosurface leading edges. I looked at
leading edges for this study, so bear in
mind that nose tips will run a little hotter still. <o:p></o:p></p><p class="MsoNormal">In this Nusselt correlation,
you evaluate boundary layer properties at the total pressure and total
temperature properties behind a normal shock at flight conditions. I used the NACA 1135 tables for this. It also uses a second viscosity evaluated at
the flight conditions. I did this for
Mach 2 to Mach 7, at the same three altitudes
as the skin panel study. The idea was to
balance convective heating against thermal re-radiation, with no conduction or active cooling, as in the skin panel study. <o:p></o:p></p><p class="MsoNormal">The results at 40 kft are given in <b>Figures 9 and 10</b>. <b>Figure 9</b> shows trends of total
temperature, and two local stagnation-region
equilibrium temperatures, one at low
emissivity, one at high emissivity. <b>Figure 10</b> superposes material service
limits on the same curves. The same data
in the same format is given in <b>Figures 11 and 12</b> at 85 kft, and <b>Figures 13 and 14</b> at 130 kft. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhOjekEpx_JtUOYleHfOhHSstOh344hSLkIhr89K6VWUmCKTMLQ2FIQfZ8GK-BdcupDJHQxHoBsLThgrrF8yesJIke42XdfeVJLxND3bvCiqrfeCwH_c4vo6XjREUTg7UGTkyyDr2yAARPdRVBxMGsK8XEoj1X1OaQP6W8p7QizzgrC7ZYinjxxTp0-SfJ7/s996/F9%20stagn%20demo%2040%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhOjekEpx_JtUOYleHfOhHSstOh344hSLkIhr89K6VWUmCKTMLQ2FIQfZ8GK-BdcupDJHQxHoBsLThgrrF8yesJIke42XdfeVJLxND3bvCiqrfeCwH_c4vo6XjREUTg7UGTkyyDr2yAARPdRVBxMGsK8XEoj1X1OaQP6W8p7QizzgrC7ZYinjxxTp0-SfJ7/w400-h236/F9%20stagn%20demo%2040%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal"><a name="_Hlk146728408">Figure 9 – Stagnation Region Soakout
Results vs Mach at 40 kft<o:p></o:p></a></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjhuyIXlPG1hxQLjY4orpL5oFrC1gn_EB8bTKIDaEGiQW5jd-8znMEYexHZS-HL0GSruqwEjrRKErRTI68BPduEe0nXjU41BndCwC8qin4EcT0U4LoqjRXobxoxNu4yf-XQEfPxm_EKU7sR35s-CatgInQduWIlmz5rwGYvLvXOTyE0m9rai2kEzuWTZ1Pj/s996/F10%20matls%20at%2040%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjhuyIXlPG1hxQLjY4orpL5oFrC1gn_EB8bTKIDaEGiQW5jd-8znMEYexHZS-HL0GSruqwEjrRKErRTI68BPduEe0nXjU41BndCwC8qin4EcT0U4LoqjRXobxoxNu4yf-XQEfPxm_EKU7sR35s-CatgInQduWIlmz5rwGYvLvXOTyE0m9rai2kEzuWTZ1Pj/w400-h236/F10%20matls%20at%2040%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 10 – Soakout at 40 kft with Service Limits, and a Speed Limit Indicated with Inconel
X-750<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiSS6_WVIrJ4_OK3TBkj6CHZFcr8fO4r49PApcwhV9vv2bCwsgwOaM-VJbB7vFey3jddtQlu2rzUjS5FFefmNg2w-VdWuLjv2at57Bj9Rvy3X2KiOHPXlSrLmAdkcEt_VUHPao3M7toP98XhC6gXAt2yZABbZlcBeNiRLZT9PsUjaOr1-5kzfocwSf8UoYC/s996/F11%20stag%2085.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiSS6_WVIrJ4_OK3TBkj6CHZFcr8fO4r49PApcwhV9vv2bCwsgwOaM-VJbB7vFey3jddtQlu2rzUjS5FFefmNg2w-VdWuLjv2at57Bj9Rvy3X2KiOHPXlSrLmAdkcEt_VUHPao3M7toP98XhC6gXAt2yZABbZlcBeNiRLZT9PsUjaOr1-5kzfocwSf8UoYC/w400-h236/F11%20stag%2085.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 11 – Stagnation Region Soakout Results vs Mach at 85
kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgzl0OVa7LrFOfXe_02JabGb5zQ0icOVdP7D-EIFj6Nxej0GLORo1L2LhDXjb5h3qWljE8ge5VvBZ3heEEBl3DHZUj_D7FiKcMTtUiXt2mgqAxTZq4wYjNkPf9Hlu2Yp4Ms-3cwS8FDaQpzHpvkMQVdA8Gv4lea6mvFYdQDn79uCJdSIy7_Rt71ISO9qLeH/s996/F12%20matls%2085kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEgzl0OVa7LrFOfXe_02JabGb5zQ0icOVdP7D-EIFj6Nxej0GLORo1L2LhDXjb5h3qWljE8ge5VvBZ3heEEBl3DHZUj_D7FiKcMTtUiXt2mgqAxTZq4wYjNkPf9Hlu2Yp4Ms-3cwS8FDaQpzHpvkMQVdA8Gv4lea6mvFYdQDn79uCJdSIy7_Rt71ISO9qLeH/w400-h236/F12%20matls%2085kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 12 – Soakout at 85 kft with Service Limits, and a Speed Limit Indicated with Inconel
X-750<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjEBud_w3dQ5S9WN6NWH-vhnhOzOF-D19GBmt6Oq2CgI13vDNiCM2LiMLa_TYtlso6GoGDQAR5vHvU6owvqj5s7PMm3wN6xRtJR_Rtd6G0f9DG87a-ee0Py2vAMWs8APdYaCd5avhr1Nyh7JwJkFVrCbICTOqLDI8MK7m_Ag7iWxo_h55udneAsXywjnuMj/s996/F13%20stag%20130%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjEBud_w3dQ5S9WN6NWH-vhnhOzOF-D19GBmt6Oq2CgI13vDNiCM2LiMLa_TYtlso6GoGDQAR5vHvU6owvqj5s7PMm3wN6xRtJR_Rtd6G0f9DG87a-ee0Py2vAMWs8APdYaCd5avhr1Nyh7JwJkFVrCbICTOqLDI8MK7m_Ag7iWxo_h55udneAsXywjnuMj/w400-h236/F13%20stag%20130%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 13 – Stagnation Region Soakout Results vs Mach at 130
kft<o:p></o:p></p><p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjN7QsGLsV2Fe8p2gpFZLQnq6ey-zDSdlrTm5uyMMFtfyNpfmarmBgS0LjuaYjr0hKA_FJy9XsgCShzTQbjnmPYkOB8zmaoT0Y7Typ8xCg06Kby2nPE-ftLHl0yWQuRAsPYW7SAZKFWJE_y2pDfMwz6vLOSmMX-IYukFEouvnDQ7FjhoegpIP9NKgClLKLy/s996/F14%20matls%20130%20kft.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjN7QsGLsV2Fe8p2gpFZLQnq6ey-zDSdlrTm5uyMMFtfyNpfmarmBgS0LjuaYjr0hKA_FJy9XsgCShzTQbjnmPYkOB8zmaoT0Y7Typ8xCg06Kby2nPE-ftLHl0yWQuRAsPYW7SAZKFWJE_y2pDfMwz6vLOSmMX-IYukFEouvnDQ7FjhoegpIP9NKgClLKLy/w400-h236/F14%20matls%20130%20kft.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 14 – Soakout at 130 kft with Service Limits, and a Speed Limit Indicated with Inconel
X-750<o:p></o:p></p><p class="MsoNormal">In <b>Figures 10,
12, and 14</b>, I have included data for the service
temperature limits and tensile strength at those limits, as part of the figure. Of the metals possibly useful for these high
speed exposures, Inconel X-750 is by far
the strongest, leading to thinner parts
of lower weight. So, I used it as the selection here, for “best” performance. Under the earlier name “Inconel-X”, this was in fact the skin material and
leading edge for the X-15 rocket plane,
which skin was a major load-bearing portion of its airframe. <o:p></o:p></p><p class="MsoNormal">Even so, the speed
limit for Inconel X-750 in a stagnation zone is only about Mach 4.9 at 40
kft, about Mach 5.2 at 85 kft, and about Mach 5.8 at 130 kft. For lateral skins, this was nearer Mach 6 at 40 kft, Mach 7 at 85 kft, and likely near or above Mach 8 at 130 kft, because the convective heat to be reradiated
is far lower for lateral skins, compared
to stagnation zones. <o:p></o:p></p><p class="MsoNormal">A good guess says the stagnation limit for Inconel X-750 is about
Mach 5.5 at 100 kft, which neatly
explains why the X-15A-2 with the drop tanks was coated all-over with an
ablative for its flights to Mach 6 and beyond,
despite the indicated survivability of its lateral skins at Mach 7+, near 100 kft. <o:p></o:p></p><p class="MsoNormal">The craft reached Mach 6.7 at 99,000 feet on flight
188, and suffered shock-impingement
heating damage to the underside of its tail,
to both lateral and stagnation surfaces.
That phenomenon drastically raises the local heating rate, but not the actual gas temperatures, as described in another of my articles on
this site: “Shock Impingement Heating Is
Very Dangerous”, posted 12 June 2017. See also NASA TM-X-1669 ““Flight Experience
With Shock Impingement and Interference Heating on the X-15-2 Research
Airplane”, dated October 1968, and written by Joe D. Watts, at the Flight Research Center, Edwards,
CA. This document is publicly
available over the internet. <o:p></o:p></p><p class="MsoNormal"><b>Stagnation Study Results:<o:p></o:p></b></p><p class="MsoNormal">Use no metals for leading edge stagnation zones that are cooled
only by re-radiation, past about Mach 5.5, and then only above 100 kft. You must instead use ablatives, or apply massive active cooling. <b>See Figure 15. <o:p></o:p></b></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEikT951ISC-SQksUF-dqELc_HNwJHRtpT4MBf0q5zDVSAKr61TskQiD-yj9AMnyxGQz5yQh0r4IYBhyyLzWfGoUQIBvM5i7oZQi29kh-u4bDTcx_FLa2W0bTRjYt5sikuHdhyJAMd-MU0QFE9Yv6XrCHXB7kle2dd8AielJ9rppP2NRuY-IauzZMs95rEDH/s996/F15%20stag%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEikT951ISC-SQksUF-dqELc_HNwJHRtpT4MBf0q5zDVSAKr61TskQiD-yj9AMnyxGQz5yQh0r4IYBhyyLzWfGoUQIBvM5i7oZQi29kh-u4bDTcx_FLa2W0bTRjYt5sikuHdhyJAMd-MU0QFE9Yv6XrCHXB7kle2dd8AielJ9rppP2NRuY-IauzZMs95rEDH/w400-h236/F15%20stag%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 15 – Results for Stagnation Zone Equilibrium <o:p></o:p></p><p class="MsoNormal">Nose tips will run slightly hotter than leading edges
(higher h values at the higher C raise Tsurf),
thereby have a somewhat lower speed limitation than leading edges. The risk with both locations is distortion
and collapse of the parts, as they
weaken rapidly with increasing overheat.
<o:p></o:p></p><p class="MsoNormal">Alloys like Rene 41 and Alloy 188 can take slightly higher
temperatures than Inconel X-750, but are
inherently weaker structurally by around a factor of 2. This is a crucial consideration, because stagnation zones see the highest
positive surface pressures on the airframe. Distorted or failed leading edges lead to
higher drag, loss of lift, and intrusion of hot gas inside the
aerosurface, something to be assiduously
avoided. In general, weaker is thicker, which is heavier. <o:p></o:p></p><p class="MsoNormal"><b>Lateral Skin Results:<o:p></o:p></b></p><p class="MsoNormal">Speed limits versus altitude for Inconel X-750 lateral skins
are about Mach 6 at 40 kft, a bit over
Mach 7 at 85 kft, and likely above Mach
8 at 130 kft. This is complicated by the
risks of shock impingement heating,
which occurrence is complex and difficult to predict, and which can do fatal damage at much lower
speeds nearer only Mach 6. <b>See Figure
16</b>. Bear in mind that the analysis
method is invalid above about Mach 7,
although the prediction is likely still crudely true. <o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjv2UKbenqCE_AndDz1WpuU0apbPUqXOg58Bt3IVum6aWUWloRkdwjLOcxvkiJpKBOi_WwPC7FLmFnFavQGr8LxJ0gr_IZlZ_RzCCHmMHLcKEcFS0ufwnYDdGwhroGQ7OkIR-9UKWWtYGSgR4uGarmxS3wxfpIRetJWKKz5UNssmbBQk2Hs3T-y2CujdQuI/s996/F16%20skin%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjv2UKbenqCE_AndDz1WpuU0apbPUqXOg58Bt3IVum6aWUWloRkdwjLOcxvkiJpKBOi_WwPC7FLmFnFavQGr8LxJ0gr_IZlZ_RzCCHmMHLcKEcFS0ufwnYDdGwhroGQ7OkIR-9UKWWtYGSgR4uGarmxS3wxfpIRetJWKKz5UNssmbBQk2Hs3T-y2CujdQuI/w400-h236/F16%20skin%20results.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 16 – Results for Lateral Skin Equilibrium<o:p></o:p></p><p class="MsoNormal">As with stagnation zones,
there are alloys that will go a little hotter, but at far lower strength. This is a crucial consideration, because in monocoque construction, the skins are an integral part of the
airframe structure, bearing much more
than just local surface pressure loads.
Weaker is thicker, which is
heavier.<o:p></o:p></p><p class="MsoNormal"><b>Remarks About Airbreathers:<o:p></o:p></b></p><p class="MsoNormal">Components associated with airbreathers (of any type) were not
studied here. The X-15 was a rocket
plane. The results above apply to both
rocket-powered hypersonic vehicles, and
to hypersonic gliders. <o:p></o:p></p><p class="MsoNormal">All airbreathers will have some sort of supersonic inlet
capture structures, some sort of
post-capture air ducting that leads to the engine device (whatever it is), and that engine device and its nozzle. The ducting,
engine device, and nozzle might
be either buried inside the airframe, or
exposed as part of the airframe. <o:p></o:p></p><p class="MsoNormal"> <b><i>Air
Inlet Components<o:p></o:p></i></b></p><p class="MsoNormal">Inlet capture features suffer worse heating effects than leading
edge (or nose tip) stagnation surfaces,
This is because they are heated (unequally) on both outside and inside
surfaces, but can re-radiate to cool
from only the exterior surfaces, with
very localized stagnation soak-out on leading edges that must stay thin and
sharp, in order to function
properly. There is little opportunity
for any conduction-as-cooling, and not
much opportunity for any active cooling.
They must also contain serious internal pressures without shape
distortion. <o:p></o:p></p><p class="MsoNormal">Buried subsonic inlet ducts will inevitably soak out to essentially
the full air recovery temperature, or
else they must be actively cooled. They cannot re-radiate, being buried inside the airframe. They must be externally insulated to protect
the rest of the airframe and its contents.
<o:p></o:p></p><p class="MsoNormal">Exposed inlet ducts are unlikely in hypersonic designs, as too much airframe drag gets added. However,
these are also internally heated,
and can only re-radiate to cool from that portion of the outside
surfaces not inside a fairing or facing the fuselage. They will still tend to approach air recovery
temperature soak-out, although not as
closely as buried ducts. <o:p></o:p></p><p class="MsoNormal"> <b><i>Combustor
and Nozzle Components<o:p></o:p></i></b></p><p class="MsoNormal">Buried or exposed combustors eventually soak out to something
in between the external and internal recovery temperatures, and will likely need active cooling. The buried combustor will take a longer time
to equilibriate, because it starts off
exposed to low airframe internal temperatures,
with a relatively low thermal conductivity for the free convection or
insulated interfaces between it and the skin.
But it will soak out very hot! <o:p></o:p></p><p class="MsoNormal">An exposed combustor can re-radiate directly to the
surroundings, while the buried combustor
cannot (while the airframe skin can), so
the exposed combustor may possibly equilibriate a little cooler than the buried
combustor. But neither has a cold “sink”
to dump heat into. They both get very
hot! <o:p></o:p></p><p class="MsoNormal">The same applies to propulsion nozzle structures, whether buried or not. <o:p></o:p></p><p class="MsoNormal"> <b><i>Turbomachinery<o:p></o:p></i></b></p><p class="MsoNormal">As for turbomachinery (compressors and turbines), these must be isolated completely from hot
intake airflow above about Mach 3 to 3.5.
Beyond that speed, the very
intake air temperature exceeds the turbine inlet temperature limits of almost
any conceivable design. The main flying
examples of these speed limitations were the XB-70 (Mach 3.0), the SR-71 (Mach 3.2), and the Mig-25 (Mach 3.5).<o:p></o:p></p><p class="MsoNormal"> <b><i>(Subsonic-Combustion)
Ramjet<o:p></o:p></i></b></p><p class="MsoNormal">Ramjet can fly faster than turbine, before hitting overheat speed limits. Flight tested but not fielded as
operational, the ASALM-PTV test vehicle
was designed to cruise steady state at Mach 4 and 80 kft, followed by an average Mach 5 terminal dive
onto its target. It did so successfully
in flight test. <o:p></o:p></p><p class="MsoNormal">In one test of ASALM-PTV,
an assembly error led to a throttle runaway incident, with the vehicle accelerating to fuel
exhaustion at Mach 6 at low altitude (near 20 kft). It suffered airframe overheat damage, but actually survived the short transient
flight and was recovered after it crashed.
<o:p></o:p></p><p class="MsoNormal">If designed for it,
ramjet could conceivably be made to work steady-state at Mach 6, or even a bit faster, perhaps.
The internal air duct and combustor/nozzle will require active cooling
for a long flight. The inlet cowl lip
surfaces will likely need to be made of a really high-melting metal, like tungsten or columbium, so that they remain both sharp and thin, without distorting. <o:p></o:p></p><p class="MsoNormal"> <b><i>Supersonic-Combustion
Ramjet (Scramjet)<o:p></o:p></i></b></p><p class="MsoNormal">Scramjet can fly faster still than ramjet, but faces similar overheat risks for its
inlet capture and supersonic isolator duct,
and its combustor and nozzle structures.
These get ridiculously difficult to design for, as speeds increase beyond Mach 7. The same can be said for airframe stagnation
surfaces and lateral skins. Short
transients and ablative materials make such flight possible, but those are neither reusable, nor are they long-range. <o:p></o:p></p><p class="MsoNormal"> <b><i>Altitude
Limits<o:p></o:p></i></b></p><p class="MsoNormal">The problem with all airbreathers, of any type whatsoever, is the “service ceiling” effect. These devices produce an altitude-dependent
characteristic trend of thrust versus speed,
with lower thrust levels in the thinner air at higher altitudes. Roughly speaking, thrust is proportional to the ambient
atmospheric pressure at altitude. So is
drag. But weight does not vary with
altitude, only with time as fuel burns
off. <o:p></o:p></p><p class="MsoNormal">The vehicle requires enough lift to offset the perpendicular
component of its weight, as it tries to
fly up an ascending path. It also
requires enough thrust to offset the sum of drag and the pathwise component of
its weight. <b>See Figure 17</b>.<o:p></o:p></p><p>
</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh0vCPw8PkvRDg6-1tKZFqDAnV6kRcLir_fY0-8QCMb3edtuqRLNHPjHrkUlOVR1o8tdpbzJtgpetmUJO0kRkBkybriRR-RuVNXXGxE-nJC2W-jA2kKldnjirAkPLCSmCIk_m_ZkmoJuNz0qt-2jfKqbfaMS6Fvd0UKDRleJIy7TAE3qzVAvSBc9aXUeL9K/s996/F17%20service%20ceiling.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="587" data-original-width="996" height="236" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEh0vCPw8PkvRDg6-1tKZFqDAnV6kRcLir_fY0-8QCMb3edtuqRLNHPjHrkUlOVR1o8tdpbzJtgpetmUJO0kRkBkybriRR-RuVNXXGxE-nJC2W-jA2kKldnjirAkPLCSmCIk_m_ZkmoJuNz0qt-2jfKqbfaMS6Fvd0UKDRleJIy7TAE3qzVAvSBc9aXUeL9K/w400-h236/F17%20service%20ceiling.png" width="400" /></a></div><p></p><p class="MsoNormal">Figure 17 – Why There Is an Altitude Limit for Airbreathers<o:p></o:p></p><p class="MsoNormal">There is an altitude at which there is insufficient thrust
to overcome drag and the weight component,
regardless of any wings that might solve the lift problem. Above that altitude, it cannot even fly level steady-state, at all.
As a rule-of-thumb at speeds in the Mach 5 to 7 range, that’s around 130 kft, almost no matter what sort of airbreather you
might design.<o:p></o:p></p><p class="MsoNormal"><b>Remarks on Active Cooling<o:p></o:p></b></p><p class="MsoNormal">This can be done reusably with a dedicated liquid
coolant, or it can be done
regeneratively with the fuel. For rocket
systems, the oxidizers are not generally
very good coolant materials, while the
fuels generally are. Either way, the coolant may not be allowed to boil inside
the cooling passages, because that leads
to vapor lock and a stoppage of coolant flow.
That in turn requires you to operate your coolant passages at very high
pressures to avoid boiling, which costs
weight, and power to run. <o:p></o:p></p><p class="MsoNormal">However, even if you
deliberately allow boiling, that reduces
heat transfer capacity of the coolant,
because the gas density is so much lower than the liquid density, for all known coolant materials. This is really a per unit volume problem, rather than a per unit mass problem, because the passage sizes are pretty much
fixed. <o:p></o:p></p><p class="MsoNormal"><b>Final Remarks<o:p></o:p></b></p><p class="MsoNormal">What I have done here is bound the problem for
rocket-propelled vehicles, or gliders, that fly hypersonically. I did this in terms of steady-state
equilibrium surface temperatures, for
lateral skins, and for stagnation zones
on nose tips and aerosurface leading edges.
<o:p></o:p></p><p class="MsoNormal">I have provided some discussions, but no numbers, for the airbreathing propulsion components
that might be applied to hypersonic vehicles.
Those are worse to thermally-manage than stagnation zones.<o:p></o:p></p><p class="MsoNormal">I have commented upon the “service ceiling” effect that
applies to any airbreather of any kind at all.
This is related to the narrow flight corridor to orbit, that resulted from the X-15 program. See also “About Hypersonic Vehicles”, posted 1 June 2022, on this site.
Plots of that corridor are in that article.<o:p></o:p></p><p class="MsoNormal">And I have commented upon the difficulties faced by any
actively-cooled designs. <o:p></o:p></p><p class="MsoNormal"><b>Note:<o:p></o:p></b></p><p class="MsoNormal">
</p><p class="MsoNormal">This article has been included in the catalog article, under the topics “aerothermo” and “ramjet”. That article is “Lists of Some Articles by
Topic Area”, posted 21 October
2021. The fastest way to reach it is to
use the navigation tool on the left side of this page. To use it,
you need the article posting date,
and its title, so in
general, jot that stuff down. Click on the year, then on the month, then on the title if more than one item was
posted that month. Simple as that. <o:p></o:p></p><p>
</p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0tag:blogger.com,1999:blog-2675974463524895416.post-48062280830019674312023-09-01T13:01:00.002-05:002023-09-18T14:53:14.387-05:00Purported SR-72 Propulsion<p>For some years now there have been marketing-hype
disclosures about Lockheed Martin’s efforts toward the “SR-72”, an intended follow-on to their famous SR-71 “Blackbird”. The hype was about hypersonic speeds above
Mach 5, and some hand-waving about an
advanced engine, usually supposedly a
combined-cycle gas turbine and scramjet (supersonic-combustion ramjet) engine.</p><p class="MsoNormal"><o:p></o:p></p>
<p class="MsoNormal">I knew the hand-waving about combined-cycle turbine-scramjet
was BS,<span style="mso-spacerun: yes;"> </span>because about the fastest
practical speed for gas turbine is about Mach 3.2 to 3.3 due to overheat damage,<span style="mso-spacerun: yes;"> </span>and about the min takeover speed for scramjet
is Mach 4.<span style="mso-spacerun: yes;"> </span>Plus,<span style="mso-spacerun: yes;"> </span>the inlet and nozzle geometries are <u>utterly
incompatible</u>.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">What that really means is that your propulsion unit has to
operate first as a gas turbine to take off and climb and accelerate to ramjet
takeover speed at about Mach 2.5,<span style="mso-spacerun: yes;"> </span>then operate
as a (<u>subsonic</u>-combustion) ramjet to accelerate above Mach 4,<span style="mso-spacerun: yes;"> </span>then finally operate as a scramjet to “fly
hypersonically” at or above Mach 5.<span style="mso-spacerun: yes;"> </span>The
ramjet and the gas turbine share similar inlet and nozzle geometries,<span style="mso-spacerun: yes;"> </span>but the scramjet is still <u>utterly
incompatible geometrically</u> with the other two.<span style="mso-spacerun: yes;"> </span>And,<span style="mso-spacerun: yes;">
</span>you must change engine type in order to slow down for a more economical
cruise.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">My suggested solution has, <span style="mso-spacerun: yes;"> </span>up to now, <span style="mso-spacerun: yes;"> </span>been “parallel-burn” propulsion:<span style="mso-spacerun: yes;"> </span><u>do not try to combine</u> the various
propulsion types into one design,<span style="mso-spacerun: yes;">
</span>instead install all 3 separately,<span style="mso-spacerun: yes;">
</span>each optimized for what it is.<span style="mso-spacerun: yes;"> </span>(Combined,<span style="mso-spacerun: yes;"> </span>it is <u>inevitable</u> that performance of
each component <u>suffers greatly</u>.)<span style="mso-spacerun: yes;"> </span>But,<span style="mso-spacerun: yes;"> </span>a major problem with parallel burn at higher
speeds (where drag is high),<span style="mso-spacerun: yes;"> </span>is that no
one of these propulsive items is a large enough fraction of the vehicle frontal
cross section area!<span style="mso-spacerun: yes;"> </span>That severely limits
the max speed attainable,<span style="mso-spacerun: yes;"> </span>likely to less
than hypersonic,<span style="mso-spacerun: yes;"> </span><u>which eliminates any
reason to have the scramjet at all</u>! <o:p></o:p></p>
<p class="MsoNormal"><b>Concept for Combining Gas Turbine with Ramjet and
Scramjet<o:p></o:p></b></p>
<p class="MsoNormal">I have since had a sort-of hybrid idea.<span style="mso-spacerun: yes;"> </span>The 3 systems can share one common supersonic
inlet capture installation,<span style="mso-spacerun: yes;"> </span>but nothing
else!<span style="mso-spacerun: yes;"> </span>The post-capture channels of the
inlet must be made variable geometry,<span style="mso-spacerun: yes;"> </span>so
that the gas turbine and the ramjet can be fed subsonic air in a diverging
channel,<span style="mso-spacerun: yes;"> </span>while the scramjet is fed
supersonic air in a constant-area channel.<span style="mso-spacerun: yes;">
</span>The supersonic channel to the scramjet <u>must</u> be “straight
through”,<span style="mso-spacerun: yes;"> </span>you <u>absolutely cannot</u>
divert a channel carrying supersonic flow,<span style="mso-spacerun: yes;">
</span>because the turn always causes shock-down to subsonic flow!<span style="mso-spacerun: yes;"> </span>Anybody who claims otherwise is spouting pure
BS!<o:p></o:p></p>
<p class="MsoNormal">The gas turbine needs to be a low-bypass ratio afterburning
design suitable for supersonic flight,<span style="mso-spacerun: yes;">
</span>and also be fitted with air bypass tubes around its core big enough so
that they can carry 100% of the air flow,<span style="mso-spacerun: yes;">
</span>tapped off <u>ahead</u> of the compressor face, <span style="mso-spacerun: yes;"> </span>and going directly to the afterburner.<span style="mso-spacerun: yes;"> </span>(In the SR-71,<span style="mso-spacerun: yes;"> </span>those engines had 25% max air bypass,<span style="mso-spacerun: yes;"> </span>tapped from the 3<sup>rd</sup> or 4<sup>th</sup>
stage of the compressor.)<span style="mso-spacerun: yes;"> </span>In that way
with 100% bypass,<span style="mso-spacerun: yes;"> </span>the afterburner can
also serve as the subsonic-combustion ramjet combustor,<span style="mso-spacerun: yes;"> </span>using the very same post-capture subsonic
inlet air channel as the turbine uses.<span style="mso-spacerun: yes;"> </span>But,<span style="mso-spacerun: yes;"> </span>we <u>do</u> need to stop the airflow into
the compressor,<span style="mso-spacerun: yes;"> </span>to avoid overheat damage!<span style="mso-spacerun: yes;"> </span><u>And</u> we need to stop backflow from the
afterburner into the turbine!<span style="mso-spacerun: yes;"> </span>Ramjet
combustor gas temperatures are <u>far higher</u> than any allowable turbine inlet
temperatures,<span style="mso-spacerun: yes;"> </span>and “leaks” lower the
ramjet pressure,<span style="mso-spacerun: yes;"> </span>lowering performance
drastically. <o:p></o:p></p>
<p class="MsoNormal">Therefore,<span style="mso-spacerun: yes;"> </span>it is a key
requirement here, <span style="mso-spacerun: yes;"> </span>when operating as a
ramjet, <span style="mso-spacerun: yes;"> </span>to stop the backflow from the
afterburner chamber from going up through the turbine into the turbine
engine.<span style="mso-spacerun: yes;"> </span>That is a serious and extremely difficult
design problem to solve!<span style="mso-spacerun: yes;"> </span>But it <u>must
be solved</u>,<span style="mso-spacerun: yes;"> </span>to prevent turbine
overheat,<span style="mso-spacerun: yes;"> </span>and to raise the achievable
chamber pressure of the ramjet, <span style="mso-spacerun: yes;"> </span>in order
to preserve its performance.<span style="mso-spacerun: yes;"> </span>Leaks are
low chamber pressure,<span style="mso-spacerun: yes;"> </span>and low pressure
is low performance.<span style="mso-spacerun: yes;"> </span>Period.<span style="mso-spacerun: yes;"> </span>That was settled long ago in tests.<o:p></o:p></p>
<p class="MsoNormal">What you “buy” with the 100% bypass and the backflow
stoppage complications, <span style="mso-spacerun: yes;"> </span>is a gas turbine
and a ramjet that share the same portion of the vehicle frontal cross
section,<span style="mso-spacerun: yes;"> </span>which then can be a <u>much larger
fraction</u> of vehicle frontal cross section,<span style="mso-spacerun: yes;">
</span>so that the top speed in ramjet can be higher,<span style="mso-spacerun: yes;"> </span>reaching the scramjet takeover range at Mach
4+.<span style="mso-spacerun: yes;"> </span><o:p></o:p></p>
<p class="MsoNormal">For scramjet takeover,<span style="mso-spacerun: yes;">
</span>you must suddenly change the inlet post-capture channel geometry to a long,<span style="mso-spacerun: yes;"> </span>straight supersonic feed to the scramjet, <span style="mso-spacerun: yes;"> </span>that is also the “isolator duct” required for
stable scramjet operation.<span style="mso-spacerun: yes;"> </span>This scramjet
must be parallel-mounted to the rest of the propulsion,<span style="mso-spacerun: yes;"> </span>and <u>must be completely separate</u>,<span style="mso-spacerun: yes;"> </span>except for sharing the supersonic capture
features.<span style="mso-spacerun: yes;"> </span>It lets you put the scramjet
on the belly of the aircraft,<span style="mso-spacerun: yes;"> </span>and to use
the vehicle aft underside as a free-expansion nozzle surface. That reduces (<u>but
does not zero</u>) the scramjet’s fraction of the vehicle frontal cross section,<span style="mso-spacerun: yes;"> </span>as opposed to that of the turbine/ramjet,<span style="mso-spacerun: yes;"> </span>to about a 50-50 split.<span style="mso-spacerun: yes;"> </span>That highly-integrated geometry in turn increases
the max scramjet speed against drag,<span style="mso-spacerun: yes;">
</span>making more-than-minimum (Mach 5) “hypersonic speed” feasible.<o:p></o:p></p>
<p class="MsoNormal">Doing these required design features is a
hellaciously-difficult problem,<span style="mso-spacerun: yes;"> </span>but does
offer a potentially-feasible solution for hypersonic flight that does <u>not</u>
involve rocket thrust to takeover speed.<span style="mso-spacerun: yes;">
</span><u>I have not even touched on the thermal management issues</u>,<span style="mso-spacerun: yes;"> </span>which may, <span style="mso-spacerun: yes;"> </span>in point of fact, <span style="mso-spacerun: yes;"> </span>be <u>fatal</u> to the concept!<span style="mso-spacerun: yes;"> </span>Suffice it to say the usual construction
techniques for the afterburner and its nozzle cannot be used,<span style="mso-spacerun: yes;"> </span>because for Mach 3.3+ speeds,<span style="mso-spacerun: yes;"> </span>there is <u>no such thing</u> as the cooling
air that those technologies require. <o:p></o:p></p>
<p class="MsoNormal"><b><i>Finally,<span style="mso-spacerun: yes;"> </span>if the
marketing hype you see does not include a propulsion system that addresses the
issues I have raised here,<span style="mso-spacerun: yes;"> </span><u>and</u> a
thermal management scheme that addresses the propulsion <u>and</u> the inlet <u>and</u>
the airframe,<span style="mso-spacerun: yes;"> </span>then I suggest that you
dismiss it as the BS that it quite evidently is!<o:p></o:p></i></b></p>
<p class="MsoNormal">A cartoon sketch of my scheme is given here as <b>Figure 1</b>.
<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhl5jqPkDxp5gExJj-zNEfEx8Dri8-4ocFgCFadVwlEgHCkmgpcuqHK4PI94k5uqa-KOdgLradH1CZ1MC1UuvG2RxC8Vo6Xaiw1jvXmIn-pW_fAgssSXN_kLnEeaZ-ZcF_m3GcFL5nXYPiCJdYzZmrgrx-CdT3m8F_MAh46MlrNeG0mucahEniUh9nbsw3v/s1017/hypersonic%20aircraft%20propulsion.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="620" data-original-width="1017" height="244" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhl5jqPkDxp5gExJj-zNEfEx8Dri8-4ocFgCFadVwlEgHCkmgpcuqHK4PI94k5uqa-KOdgLradH1CZ1MC1UuvG2RxC8Vo6Xaiw1jvXmIn-pW_fAgssSXN_kLnEeaZ-ZcF_m3GcFL5nXYPiCJdYzZmrgrx-CdT3m8F_MAh46MlrNeG0mucahEniUh9nbsw3v/w400-h244/hypersonic%20aircraft%20propulsion.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – A Possible Means to Combine Gas Turbine Takeoff and
Landing with Scramjet Dash<o:p></o:p></p>
<p class="MsoNormal"><b>Rocket-Boosted Ramjet Is a Much Better Way<o:p></o:p></b></p>
<p class="MsoNormal">Actually, I still
prefer my parallel-burn, completely
separate, rocket and ramjet solution, and just forget the scramjet! To take off,
climb, and accelerate to around
Mach 2.5 does not require all that big a rocket engine, or all that much propellant. The subsonic-combustion ramjet takes over at
about Mach 2.5, and supports supersonic
cruise <u>much more economically</u> in the vicinity of Mach 3, but with enough frontal cross section
fraction to support supersonic dash speeds to Mach 5, or possibly even Mach 6. And that <u>is</u> hypersonic! No scramjet required! It just has lower specific impulse at
hypersonic speeds, <u>as does the scramjet</u>.
However,
you do <u>not</u> have to change propulsion to slow down to cruise!<o:p></o:p></p>
<p class="MsoNormal">If you include some small liquid rocket propulsion, your landing is <u>not entirely “dead-stick</u>”. Just fire up the liquid rockets to divert or
go-around. I find that to be a far safer
and more practical solution, manned or
unmanned! <o:p></o:p></p>
<p class="MsoNormal">The main mass of booster propellant to reach ramjet
takeover, is likely a solid packaged
within the ramjet combustor as an “integral rocket ramjet” booster (IRR booster). There are two reasons for this: (1) the booster needs to be big to have the very
high thrust to accelerate very quickly to ramjet speed, to reduce the aerodynamic drag losses to
tolerable values, and (2) there are no
air-cooled technologies available for the combustor and nozzle internal heat
protection at these flight speeds, since
there is no such thing as “cooling air” above about Mach 3.2 to 3.3; thus the only technological solutions for
combustor and nozzle are one-shot ablatives.
The IRR is proven, existing
1-shot missile technology. <o:p></o:p></p>
<p class="MsoNormal">That last says you need to pull the entire ramjet combustor unit
out, and replace it, after every flight! It therefore might as well contain an
integral solid booster, just like what
has proved so successful in missile work. You need the big boost to ramjet speed <u>only
once per mission</u>! The smaller liquid
rockets let you fly the plane at speeds below ramjet speed, for the approach and landing. <o:p></o:p></p>
<p class="MsoNormal">See <b>Figure 2</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_Lr345RnBnG5O3bfpB3boWmM1xhaavWCP-xoyvqCr6521V4-ptyN2gePoZXvLsQU79GbuGQepZQVnWyhHaTPI0H8CPloJRH3i8bICW1AgQDqRO1eGNTX_iBuEpfI4qRw7HEJIii3oNIYhvkN65Tn4mElm-lviDoPqModdtI8RfikIjQyGmOj0g1pZmDoV/s985/ramjet%20plane.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="594" data-original-width="985" height="241" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg_Lr345RnBnG5O3bfpB3boWmM1xhaavWCP-xoyvqCr6521V4-ptyN2gePoZXvLsQU79GbuGQepZQVnWyhHaTPI0H8CPloJRH3i8bICW1AgQDqRO1eGNTX_iBuEpfI4qRw7HEJIii3oNIYhvkN65Tn4mElm-lviDoPqModdtI8RfikIjQyGmOj0g1pZmDoV/w400-h241/ramjet%20plane.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – Rocket-Boosted Ramjet as a Means to Achieve
Hypersonic Dash<o:p></o:p></p>
<p class="MsoNormal"><b>Figure 3</b> shows some details about how the
cartridge-loaded ramjet combustor and nozzle is also its own integral rocket ramjet (IRR) booster. The craft need accelerate only once to ramjet
takeover speed, and the IRR booster does
that job, then transitions to ramjet
thrust in about 0.1 sec (as demonstrated by ASALM-PTV in flight). The liquid rockets are much smaller, and mainly serve to keep the descent and
landing from being totally “dead stick” (with no go-around or divert capability).
<o:p></o:p></p>
<p class="MsoNormal">Combustor and nozzle heat protection is by ablative
materials, which cannot be re-used. So, the
IRR unit <u>must be replaced</u> for every flight. In this concept, there must be airframe structure to support
the vertical tail, so the IRR unit
resides inside this airframe, not
exposed to hypersonic external aeroheating.
That greatly simplifies the thermal management, to something the ablatives can easily handle
for very long burns. The case can be
power-washed out, refitted with
ablatives, and cast with another
propellant charge. On-pavement recovery
has little in the way of risk to support this kind of reuse. <o:p></o:p></p>
<p class="MsoNormal">By making the bottom flat with the bifurcated inlet
ducts, there is little need for wing
area in supersonic flight above about Mach 3,
but there is room for the small liquid rockets aft of the inlets ducts! The wing is really sized for a tolerable
landing speed, with the delta planform
allowing high angle of attack without stalling.
It is mostly just parasite drag at high speeds, so there are many design tradeoffs here. However,
at very high altitudes in very thin air,
the wing allows sufficient lift generation at lower angles of attack
that correspond to lower drag-induced-by-lift.
This may help extend cruise range,
and certainly might help extend the service ceiling. The “right” wing is quite likely smaller
than the one sketched on the figure. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhZeehL0c_fHSteztYYnby8uYotdN-JirxUa93Z7d9JAO6Cxmjb-ZoyuXxoEtyhkFnhfpBdry2yP0Q3lMeb7kMTG-OfOXqYMIVmP3G89mJnJ0OpXprVEo-e7Q4Guu0hrLdyJZ6YWTTp-NkHdjPOxaIf1mSBOAFm82J9p6ZMKbkHtEhbgd6A1-OOCuImGxMu/s985/cartridge%20load%20IRR%20engine.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="594" data-original-width="985" height="241" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhZeehL0c_fHSteztYYnby8uYotdN-JirxUa93Z7d9JAO6Cxmjb-ZoyuXxoEtyhkFnhfpBdry2yP0Q3lMeb7kMTG-OfOXqYMIVmP3G89mJnJ0OpXprVEo-e7Q4Guu0hrLdyJZ6YWTTp-NkHdjPOxaIf1mSBOAFm82J9p6ZMKbkHtEhbgd6A1-OOCuImGxMu/w400-h241/cartridge%20load%20IRR%20engine.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – Cartridge-Loaded Ramjet Combustor with IRR
Booster<o:p></o:p></p>
<p class="MsoNormal">In cruise at about Mach 3,
the ramjet specific impulse (Isp) should be in the neighborhood of 1000-1300
secs. Running richer at full ramjet
thrust for Mach 5+ dash, the ramjet Isp
is likely nearer only 700-800 sec. The
liquid rockets are lower-pressure units that are simply pressure-fed the LOX, and little bit of the same thermally-stable
kerosene that the ramjet uses. It would be
realistic to expect about 300 sec of Isp out of them. The solid booster, at about 85-87% solids, would achieve a sea level Isp near 250-255
sec. <o:p></o:p></p>
<p class="MsoNormal">This plane could actually take off using the small
rockets, like the “rocket racer”
did, although zero-length launch from a
ramp is also very feasible, since the
integral rocket booster accelerates the airplane at 5+ gees. Once leaving the pattern, you pull up sharply, fire up the solid booster and shut down the
small rockets. Seconds later, you do ramjet takeover at about Mach 2.5 while
climbing very steeply, and at much
higher altitude. The ramjet then takes
you to cruise conditions, and also hypersonic
dash. <o:p></o:p></p>
<p class="MsoNormal">At mission’s end, you
start your approach in ramjet, but shut
it down as you decelerate below Mach 2.5,
making most of the rest of the approach in glide. As you near the field, use the small liquid rockets as necessary to
divert or to go around for a missed approach.
There is only one boost to ramjet takeover per mission, but the small rockets can be used multiple
times for multiple purposes in a mission.
<o:p></o:p></p>
<p class="MsoNormal">You swap out the spent combustor unit for a fresh one, and refill the kerosene and oxygen
tanks. With on-ramp recovery, spent combustor refurbishment is also a very
low risk possibility. Easy!<o:p></o:p></p>
<p class="MsoNormal">None of these considerable existing-technology advantages
obtain with the sort-of combined-cycle gas turbine/ramjet/scramjet craft described
above. There are still
missing-technology items with it, but
not with this rocket-ramjet airplane. <o:p></o:p></p>
<p class="MsoNormal"><b>Related Information:<o:p></o:p></b></p>
<p class="MsoNormal">If you want to see more about how supersonic inlets really
work, and how they are adapted to ramjet
versus gas turbine, please see on this
site “Fundamentals of Inlets”, posted 9
November 2020. <o:p></o:p></p>
<p class="MsoNormal">If you want to see more about how (subsonic combustion)
ramjets really work, please see “How
Ramjets Work”, posted 1 December 2022, and “Primer On Ramjets”, posted 10 December 2016. <o:p></o:p></p>
<p class="MsoNormal">The general issues that must be addressed for hypersonic
vehicles are discussed in “About Hypersonic Vehicles”, posted 1 June 2022. A peculiar problem with high hypersonic flight
is discussed in “Plasma Sheath Effects in High Hypersonic Flight”, posted 18 September 2022, which debunks some of the widely-circulating
myths about “unstoppable” hypersonic missile weapons. <o:p></o:p></p>
<p class="MsoNormal">If you want to see what an integral solid booster is, please see “Solid Rocket Analysis”, posted 16 February 2020, and concentrate on the low L/D keyhole slot
grain design therein. How the internal
ballistics of solid propellant devices work is well-explained. There is also information on achievable burn
rates, and on safety sensitivity data. <o:p></o:p></p>
<p class="MsoNormal">The thermal management issues are discussed in more detail
in “On High-Speed Aerodynamics and Heat Transfer”, posted 2 January 2020, “Heat Protection is the Key to Hypersonic
Flight”, posted 4 July 2017, and “Shock Impingement Heating Is Very
Dangerous”, posted 12 June 2017. <o:p></o:p></p>
<p class="MsoNormal">Flameholding in the ramjet wasn’t an issue discussed
here, but if you are interested, that is discussed in “Ramjet
Flameholding”, posted 3 March 2020. Something similar applies to scramjet, and something somewhat different (but still similar)
applies to gas turbine can combustors. That
article makes clear why the usual V-gutter and can stabilizers cannot work at
speeds past about Mach 3.3, and what
will work.<o:p></o:p></p>
<p class="MsoNormal">There is a whole catalog article, sorted by topic area, of many of my technical articles posted on
this site. It is “Lists of Some Articles
By Topic Area”, posted 21 October 2021. There is some duplication from list to
list, where the topic areas
overlap. It does have topic areas for
ramjet, for rocket stuff, and for high-speed aero-thermo-dynamics and
heat transfer. I do try to keep that article
updated and current. <o:p></o:p></p>
<p class="MsoNormal">You can use the navigation tool on the left side of this
page to access any of these articles very quickly. Just jot down the titles and dates. Then click on the year, the month,
and finally the title if more than one was posted that month. <o:p></o:p></p>
<p class="MsoNormal"><b>One Final Note:<o:p></o:p></b></p>
<p class="MsoNormal"><u>All of this was done with open sources</u>! I have seen no classified information for
nearly 3 decades now, since I last held
a clearance and had a need-to-know. But
it is quite likely that any “real” SR-72 vehicle will be considered a
classified design by the government,
much as the SR-71 was. About 4
decades ago, I roughed-out a vehicle somewhat
similar to the rocket-ramjet hypersonic craft outlined here, from only open sources. (If you really know what you are doing, open sources are all you need.) That design concept was confiscated by the
FBI and classified by the Pentagon. They
were exploring SR-71 replacements, even
way back then. If this current one
disappears off my site, then it happened
again.<o:p></o:p></p><p class="MsoNormal">--------------</p><p class="MsoNormal"><b><u>Update 9-5-2023</u>:</b> I took some time to rough-out the characteristics of a <u>rocket ramjet airplane design</u>, and along the way found a major choice to be made. Since this was not already done in the original article, see first the intended flight profiles, <b>given in Figure 4</b>. The plane could take off from a runway using its small-rocket power, leading to the big solid booster ignition away from the airport, or it could be launched zero-length from an inclined ramp, directly with the big booster. Climb and acceleration to cruise speed (and to dash speed) is by the ramjet. Most of the approach to landing is “dead stick” glide, but with the small liquid rockets available, to divert, or to go around for a missed approach.</p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj78PkszYM3oMVEsXUMYDZHdP0Q_ypOg91cukjBB514Tv_kLDJJJZylnuI3Nw2HrVeI7Pu7z_wpWgC3Fsm_miEyZpnhkHvgNpYygUph-tIM1SXmBEHdoMuM1BWmQxmNVlKLGLQ3elUe02YG4-S7rUqpgXl9WvLBzKuu0hRG10BWTQvAg5MXq4C7qU_yF0Vh/s985/F4%20flight%20profiles.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="594" data-original-width="985" height="241" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj78PkszYM3oMVEsXUMYDZHdP0Q_ypOg91cukjBB514Tv_kLDJJJZylnuI3Nw2HrVeI7Pu7z_wpWgC3Fsm_miEyZpnhkHvgNpYygUph-tIM1SXmBEHdoMuM1BWmQxmNVlKLGLQ3elUe02YG4-S7rUqpgXl9WvLBzKuu0hRG10BWTQvAg5MXq4C7qU_yF0Vh/w400-h241/F4%20flight%20profiles.png" width="400" /></a></div><p class="MsoNormal">Figure 4 – Concept Flight Profiles<o:p></o:p></p>
<p class="MsoNormal">I literally sized a paper liquid rocket design that uses LOX
and the ramjet fuel (thermally-stable kerosene), but is a very simple pressure-fed
system. The design goal here was <u>simplicity
above all else</u>, so that reliability
would be highest. This kind of thing
should be utterly trouble-free, at the
cost of somewhat lower performance. I
did not choose a specific igniter, but I
did indicate that the igniter is linked to the on-off valves for the
propellants. It fires when they
flow, for some small set time interval. <o:p></o:p></p>
<p class="MsoNormal">The pressurant for the propellant is dry nitrogen, commonly available in 2200 psig bottles. It is likely an airframe-mounted vessel that is
filled on the apron from standard gas bottles.
The regulators are set to deliver 700 psig to the propellant tanks, so that a bit over half of the gas vessel
pressure drop is available during the mission.
Assuming the pressure drop through the passages and injector plates is
about 200 psi, a max chamber pressure of
500 psia seems reasonable. 2:1 pressure
turndown is easily achieved. <o:p></o:p></p>
<p class="MsoNormal">The 15 degree conical nozzle is designed for expansion to
11.2 psia, so that the expected
separation backpressure at half pressure is still very slightly above sea level
atmospheric. That way, nozzle flow separation is never a concern! Expected performance data is <b>shown in
Figure 5</b>, including the small-rocket
frontal thrust density value, based on
its exit area.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj574WIyFPtFUNvSaHGCneaMykpDLk4c9S_cPDUg4VweQhkWVAKGWNGuVVipQZym-VZZL7PJsx1XNMjz8NoJr2ua1u8mzd3gQwd_tYo6S7uWd4Ejdf0oDMWjGwF8Ixr6qLPGXlhgCTCToux_qs0BYvO-V_8JiD-KF9sRM_kPxQNZ3Nc0yyssXGNfIUnUmQ9/s985/F5%20small%20rockets.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="594" data-original-width="985" height="241" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEj574WIyFPtFUNvSaHGCneaMykpDLk4c9S_cPDUg4VweQhkWVAKGWNGuVVipQZym-VZZL7PJsx1XNMjz8NoJr2ua1u8mzd3gQwd_tYo6S7uWd4Ejdf0oDMWjGwF8Ixr6qLPGXlhgCTCToux_qs0BYvO-V_8JiD-KF9sRM_kPxQNZ3Nc0yyssXGNfIUnUmQ9/w400-h241/F5%20small%20rockets.png" width="400" /></a></div><p class="MsoNormal">Figure 5 – Roughing Out a Small Liquid Rocket System
Emphasizing Simplicity Above All<o:p></o:p></p>
<p class="MsoNormal">I had some old ramjet data predicted for a design with inlet
shock-on-lip Mach number 2.5, using
kerosene fuel at equivalence ratio ER = 1.10 for max thrust without excessive
waste. These data were for Mach numbers from 2 to 6 at 40,000 feet (40 kft) on
a US 1962 standard day. I curve-fit the
variations in thrust and specific impulse vs Mach number at 40 kft, and recorded the key area ratios and size of
the sized engine. I had no data at sea
level or at 85 kft, but instead just
ratioed the thrusts by the ratio of atmospheric pressures. That is not “right”, but it is pretty close. It was easy to divide the installed ramjet
thrust by its nozzle exit area, to get
the frontal thrust density for the design study. I took an educated guess for the leaned-back
cruise specific impulse at Mach 3 cruise,
at 85 kft.<o:p></o:p></p>
<p class="MsoNormal">I also had some old vehicle drag data based on information
from Hoerner’s old “drag bible”. It
includes nose pressure drag, lateral
skin drag, aerosurface drag, and base drag effects. It is uncorrected for the drag area
reductions associated with the chin inlet mounting, and for the propulsion plumes coming from the
base. That makes these drag values a
probable over-estimate by a few-to-several percent, but at least the trend with Mach number is
correct.<o:p></o:p></p>
<p class="MsoNormal">The drag and ramjet thrust density and specific impulse data
are <b>given in Figure 6</b>. <o:p></o:p></p>
<p class="MsoNormal">I had an old IRR booster grain design in my records. It is for the wrong size, but the L/D proportion is not too far
wrong. It was easy to compute its thrust
per unit exit area, for a scaleable
frontal thrust density F/Ae = 18,350 psf to use in this study. The detail internal ballistics are not quite
right, but the frontal thrust density is
in the ballpark, regardless. Some selected data are <b>shown in Figure 7</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiMwchu6_5Zu_VWACYcxMP1o3GjTAGv3m1npEpYHXkUZx_mHQLH1MSiiGO8EIeX_sqFcXTH4CGnBxjHnQaje28QegV9OHwH4pzT6FAr8jH6AsLuIu13VL9FZkksZbyfV5ndLJzsFw2zNIYcDzYO4PLEgUesUOyNOnqgTAwyC2jdvW5i-mu2eyGN0B_yzfwV/s1019/F6%20RJ%20source%20data.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="531" data-original-width="1019" height="209" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEiMwchu6_5Zu_VWACYcxMP1o3GjTAGv3m1npEpYHXkUZx_mHQLH1MSiiGO8EIeX_sqFcXTH4CGnBxjHnQaje28QegV9OHwH4pzT6FAr8jH6AsLuIu13VL9FZkksZbyfV5ndLJzsFw2zNIYcDzYO4PLEgUesUOyNOnqgTAwyC2jdvW5i-mu2eyGN0B_yzfwV/w400-h209/F6%20RJ%20source%20data.png" width="400" /></a></div><p class="MsoNormal">Figure 6 – Rescaling Ramjet Performance From Some Old, Limited Data<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjrg0xv1w4ob6li5jC_H8EzGP4Lnl6oe6fVbzNQlNKA2U55qAVezJtpaDfmdt75Vr8xra16jNLH7wJ_ZcUs6fQ50zd3nyiksfVFVJKF0SmWdWSkKhWmqzxaSAPDV7nn_XYqM7Aiceyy9qvkfua94mS752r8YCvjfjg55wiFjekKbL1KsWqvicjKrQlAak5p/s1013/F7%20IRR%20source.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="607" data-original-width="1013" height="240" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjrg0xv1w4ob6li5jC_H8EzGP4Lnl6oe6fVbzNQlNKA2U55qAVezJtpaDfmdt75Vr8xra16jNLH7wJ_ZcUs6fQ50zd3nyiksfVFVJKF0SmWdWSkKhWmqzxaSAPDV7nn_XYqM7Aiceyy9qvkfua94mS752r8YCvjfjg55wiFjekKbL1KsWqvicjKrQlAak5p/w400-h240/F7%20IRR%20source.png" width="400" /></a></div><p class="MsoNormal">Figure 7 – An Older Grain Design Used To Rescale IRR Booster
Performance<o:p></o:p></p>
<p class="MsoNormal">The original notion of the flat-bottomed airframe with the
bifurcated inlet, finally sized-out
capable of reaching Mach 5, with the ramjet
exit area A6 proportion to the vehicle frontal blockage area Sx reaching A6/Sx
= 0.623, <b>as indicated in Figure 8</b>. This is less ramjet frontal thrust density
than originally desired, <u>which is
what limited the max dash speed to Mach 5</u>.
The data include a preliminary weight statement and some estimated
component lengths. Gross cruise range
exceeds 3000 nmi, at 85 kft. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhOfw04hgY0j01XHzzDz-B_0yHSvc9zQlVmcTUOj1JV_B4y0dZbR_7lIA_Ge12UPMpLmu8XZvFLC1EPvSA6uzL2HsGaXcbq0w6igBFrMF9UfLxVi29RctbLBxes5coegRGgmfHejynIjD2iBkZnbb75-bB50jQt6P-LC3r0AgIz5QKUE2GECtZamNQeg9fF/s1303/F8%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="613" data-original-width="1303" height="189" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhOfw04hgY0j01XHzzDz-B_0yHSvc9zQlVmcTUOj1JV_B4y0dZbR_7lIA_Ge12UPMpLmu8XZvFLC1EPvSA6uzL2HsGaXcbq0w6igBFrMF9UfLxVi29RctbLBxes5coegRGgmfHejynIjD2iBkZnbb75-bB50jQt6P-LC3r0AgIz5QKUE2GECtZamNQeg9fF/w400-h189/F8%20results.png" width="400" /></a></div><p class="MsoNormal">Figure 8 – Results for the Flat-Bottomed Airframe With
Bifurcated Inlet<o:p></o:p></p>
<p class="MsoNormal">The radial distance from vehicle outer mold line to the case
or fuel tank OD is a critical variable,
as well. <u>There must be some
such distance</u>, to isolate thermally
the hot lateral skins from the vessels containing fuel or solid
propellant. That would include some
high-temperature mineral wool insulation.
<o:p></o:p></p>
<p class="MsoNormal">Initially I set this at 6 inches, and could not exceed Mach 4. Setting it to 3 inches got me not quite to
Mach 5. Resetting it to 2 inches actually
got me to Mach 5. <u>But that is about
all I can realistically squeeze out of this design concept</u>! The strakes containing the bifurcated inlet
and small liquid rocket equipment are just too large, driven by the required air inlet duct branch
sizes. <o:p></o:p></p>
<p class="MsoNormal"><b><i>That trend illustrates the crucial role frontal thrust
density plays in high supersonic, low
hypersonic flight. There is <u>no</u>
getting around this, it is quite fundamental.
<o:p></o:p></i></b></p>
<p class="MsoNormal">An alternative design concept would not bifurcate the
inlet. Instead it would pass through the
fuel tank on its way to the engine,
within an airframe of round cross section. That makes the tank longer. There would be no plenum, but there would need to be a space in which
to S-duct the inlet from the bottom up to the central axis. The wing would have to move up to a mid-wing
mount, likely just a double delta
planform. The small rocket system would
have to be mounted in the base of the vertical tail fin, much like the one used in the NF-104
design. <o:p></o:p></p>
<p class="MsoNormal">I re-ran this alternate configuration, getting the results <b>shown in Figure 9</b>. The top dash speed reached Mach 5.5, reflecting the much larger ramjet frontal
thrust density associated with A6/Sx = 0.844.
It packages less fuel mass, but
it also has less cross section area producing drag, so the drag (and thrust requirement) is
lower. The gross cruise range figure is then
just about the same, as a result. <o:p></o:p></p>
<p class="MsoNormal">If the ramjet propulsion were exposed at the rear, being all of the aft airframe cross
section, A6/Sx would be a bit higher
still (very nearly 1.0), and the top
dash speed would then approach Mach 6,
the same way it did with ASALM-PTV on the one flight test in 1980. But such exposed propulsion is a much tougher
thermal problem to solve for long-duration burns.<o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEinUQeTa38PLeAaMBZHwkaXqgtc_rqfgUcnfp2JyRaxcjTtOok4_ShJbGKPzUP-p0dOWCOjXm5OAuEMbGvfOiT7BE-TtSlNCGkCx1rJ30uYmV5dlxVAoa7eBtgfm6rcqzRee3rEzpxaX_QZqVlkaGz_uPZS4_2g34A0nF2p3CLkWd9BmQrMKGk0GzoP9dQH/s1331/F9%20alt%20results.png" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="628" data-original-width="1331" height="189" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEinUQeTa38PLeAaMBZHwkaXqgtc_rqfgUcnfp2JyRaxcjTtOok4_ShJbGKPzUP-p0dOWCOjXm5OAuEMbGvfOiT7BE-TtSlNCGkCx1rJ30uYmV5dlxVAoa7eBtgfm6rcqzRee3rEzpxaX_QZqVlkaGz_uPZS4_2g34A0nF2p3CLkWd9BmQrMKGk0GzoP9dQH/w400-h189/F9%20alt%20results.png" width="400" /></a></div><p class="MsoNormal">Figure 9 – Results for the Round Section Mid-Wing Airframe
with Inlet Through Fuel Tank<o:p></o:p></p>
<p class="MsoNormal">Bear in mind that all of these are crude estimates, only within about 10%, at best.
However, that is good enough to
determine that dash speed nearer Mach 6 will trade off against the far-more
severe thermal management problems with exposed propulsion. Meanwhile,
if Mach 5 dash is “good enough”,
the flat-bottomed low-wing airframe with the bifurcated inlet is quite
feasible. <o:p></o:p></p>
<p class="MsoNormal">Or if Mach 5.5 dash is absolutely required, the better choice is the round airframe with
center-duct inlet and a mid-mounted wing.
That one will be somewhat more challenging to detail-design, and it will have less volume available within
its nose. (You get what you pay for.) <o:p></o:p></p>
<p class="MsoNormal">Also bear in mind that the <u>next most important
feasibility item is thermal management</u>.
Those calculations have yet to be explored. <o:p></o:p></p><p class="MsoNormal"></p><p class="MsoNormal">---------------</p><p class="MsoNormal"><b><u>Update 9-18-2023</u>:</b> As it says in the previous update, the thermal management issues still need
exploration, in order to determine feasibility. Here is an initial exploratory look. <o:p></o:p></p><p class="MsoNormal">
</p><p class="MsoNormal">First, look at “typical”
lateral skins parallel to the oncoming stream.
These could be on aerosurfaces away from leading edges, or on fuselages away from nose tips and inlet
capture features. This uses a flat plate
convection model that accounts for both compressibility and the effects of
viscous dissipation. Overall setup and
results are <b>given in Figure 1</b>. Conditions
at the edge of the boundary layer would not be very far from freestream conditions, not enough to make a great deal of difference
in the film coefficient, so this
analysis just uses free stream. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjvxZdHnx8h4pcZZaj3jrM9foE68TH02caqpAZnQChbWhSy4qeNAHTFHiCv1yTFqmbXzZBT9gcq0yi0u0vzcnaKmX8nY4UnIflNyu8fJlcHKwC36NF3s0-PSf095fd6BvGll61NDEqYojaF48xJhgW1PDqwbgWADy_OGTTdGs6oeX6vSIe5DwJaU1-dkREr/s996/some%20high%20speed%20airframe%20exposures.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="548" data-original-width="996" height="220" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEjvxZdHnx8h4pcZZaj3jrM9foE68TH02caqpAZnQChbWhSy4qeNAHTFHiCv1yTFqmbXzZBT9gcq0yi0u0vzcnaKmX8nY4UnIflNyu8fJlcHKwC36NF3s0-PSf095fd6BvGll61NDEqYojaF48xJhgW1PDqwbgWADy_OGTTdGs6oeX6vSIe5DwJaU1-dkREr/w400-h220/some%20high%20speed%20airframe%20exposures.png" width="400" /></a></div><p class="MsoNormal">Figure 1 – Thermal Analysis of a “Typical” Lateral Skin
Panel<o:p></o:p></p>
<p class="MsoNormal">In the figure are plotted total temperature Tt, recovery temperature Trec, two curves representing equilibrium panel
temperatures, and the recommended max
service levels for several possible panel materials. <o:p></o:p></p>
<p class="MsoNormal">The analysis included not only convection to the panel, but also thermal re-radiation from the
panel, as its primary method of
cooling. This was done for a typical low
emissivity, and a typical high emissivity. Also included were two paths for minor cooling
effects due to conduction into the interior.
One was through a low density mineral wool insulation layer, occupying nearly the same area as the panel. The other was through a minor area
representing the conduction path through whatever structures attach the skin
panel to the rest of the airframe,
presumed metallic, and of a
length comparable to the insulation thickness. <o:p></o:p></p>
<p class="MsoNormal">For reference, a
completely uncooled panel would soak out to the recovery temperature. At speeds under roughly Mach 4, the panel’s surface thermal emissivity does
not make much difference, since the
temperatures are low enough that there is not much thermal re-radiation. However,
above Mach 4, the panel
emissivity makes a great deal of difference,
with high emissivity (dull black surface) much better.<o:p></o:p></p>
<p class="MsoNormal">Note how organic composite panels are no good above (at most)
Mach 2, and that presumes adequate
strength at the max temperature of about 200 F,
which presumption is seriously in question. Aluminum is useless above about Mach
2.5, which explains very well why most
fighters made of it, have max dash
speeds of only just about Mach 2.5. <o:p></o:p></p>
<p class="MsoNormal">A lot of folks think titanium is a high temperature
material, but that is mistaken. Its max service temperature is 600 to 800 F
(800 F shown), which is good to a most
about Mach 3.5-ish, presuming a
highly-emissive surface. That explains
very neatly the max flight speeds of about Mach 3.2 for the SR-71, which had a dull black finish. <o:p></o:p></p>
<p class="MsoNormal">Above 1500 F capability,
there are only some stainless steels,
and 3 exotic alloys that are not steels.
Of these, only one has truly high
temperature capability at 1800 F <u>plus</u> high tensile strength: Inconel X-750 (formerly simply known as “Inconel-X”). Which neatly explains the choice of “Inconel-X”
skins on the X-15 rocket plane. The
difference between the low and high emissivity effects is the difference of
about a full Mach number for survival of lateral skins at full strength: Mach 6 if high emissivity, only Mach 5 if low. Which in turn neatly explains why the X-15 had
a dull black finish. <o:p></o:p></p>
<p class="MsoNormal">Thermal analysis of nose tips and leading edge pieces is
much harder to approximate with these simple by-hand techniques. The actual stagnation zone seeing full
stagnation heating is quite small. The large
lateral areas also see convection approximatable with the flat plate model, but at edge of boundary layer conditions
crudely approximated as those behind the oblique shock corresponding to a 10
degree flow deflection. There is thermal
re-radiation cooling from both the stagnation zone, and the lateral surfaces. There is even conduction cooling through the
thickness of the part, moving toward
where it attaches to the rest of the structure. This concept is <b>illustrated
in Figure 2</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg4N-UmkFWUpYkcisvVyz6Uff5a28DhtTA3wkKMaYa9z8AZEHKUbLhLyuScUbdV-CGAxwmnt2bk0JTesoQxEvUmgczY2KTxKE4qokJgWO4DwzupYEdPiYo2dJ1KzfXyNYi3D1aSGm4nneNglp7AoE_6Z4IcGsEuAlomwr_MJGVf7neP1FL1R2JTQKxCPiHI/s996/stagnation%20pieces%20-%20Copy.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="548" data-original-width="996" height="220" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg4N-UmkFWUpYkcisvVyz6Uff5a28DhtTA3wkKMaYa9z8AZEHKUbLhLyuScUbdV-CGAxwmnt2bk0JTesoQxEvUmgczY2KTxKE4qokJgWO4DwzupYEdPiYo2dJ1KzfXyNYi3D1aSGm4nneNglp7AoE_6Z4IcGsEuAlomwr_MJGVf7neP1FL1R2JTQKxCPiHI/w400-h220/stagnation%20pieces%20-%20Copy.png" width="400" /></a></div><p class="MsoNormal">Figure 2 – “Typical” Thermal Equilibrium Considerations for
a Leading Edge Piece<o:p></o:p></p>
<p class="MsoNormal">The results did not validate the equilibrium model. In all cases attempted, the convection into the lateral surfaces
(both top and bottom together) simply overwhelmed the effects of stagnation
heating convection, and also the numbers
for all three of the cooling paths. The “equilibrium”
temperatures to balance the mathematical model were above the oncoming stream
total temperature, which is the maximum
soak-out temperature the part could see.
<u>We must therefore conclude that in the absence of active cooling
means, these leading edge parts will
rather quickly soak out to the oncoming stream total temperature, or very near to it</u>. <b>See Figure 3</b>. <o:p></o:p></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg0HlySWCj1QhqHeJXMFT4tM4MO1fqAD7LGqZbZvWgAQDo0CDBUiInfQ4ydDbS6guU_2Y5hwtdEiBo7NZs5iSv1fowvRdHLAkl_gGZS13ENAxz75R9ieXOEd9mGvGPuSumHGVP17gm_Cf1Izovb1aKFskZjYX8XSAcFuLjHJvqDxrGtKXyTEuYl-jCS44q2/s1044/stagn%20results.png" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="589" data-original-width="1044" height="226" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEg0HlySWCj1QhqHeJXMFT4tM4MO1fqAD7LGqZbZvWgAQDo0CDBUiInfQ4ydDbS6guU_2Y5hwtdEiBo7NZs5iSv1fowvRdHLAkl_gGZS13ENAxz75R9ieXOEd9mGvGPuSumHGVP17gm_Cf1Izovb1aKFskZjYX8XSAcFuLjHJvqDxrGtKXyTEuYl-jCS44q2/w400-h226/stagn%20results.png" width="400" /></a></div><p class="MsoNormal">Figure 3 – Leading Edge Piece Results<o:p></o:p></p>
<p class="MsoNormal">The Inconel-X material as a leading edge piece may or may
not need its full strength to withstand the local wind pressures upon it. Roughly speaking, it reaches its max service temperature
limit, or a bit above, at about Mach 5. Mach 6 is very near the melting point for the
material. This very neatly explains why the X-15A-2 vehicle was coated with a pink
silicone rubber ablative and white ceramic paint topcoat, for high-speed flights past Mach 5. On flight 188, with Pete Knight flying it, it reached Mach 6.7. There was extensive airframe damage from
simple overheat in multiple stagnation regions,
and near-fatal shock-impingement heating underneath the tail
section. <o:p></o:p></p>
<p class="MsoNormal">What that really tells us is that for long flights beyond
Mach 5, one must either do
high-capability active cooling, or else use
ablative materials for the leading edges and nose tips. Active cooling will be very heavy, and very expensive in terms of the power to
run it. Ablatives will require
replacement, at worst after every
flight, or at best after every few
flights. The ablative approach is
exactly what was done with the Space Shuttle and its derivative the X-37B, and also the old X-20 design never built. <o:p></o:p></p>
<p class="MsoNormal">Remember: if you have
airbreathing propulsion, the inlet
capture features are even more challenging than leading edges and nose tips, and the buried ducts simply <u>will require</u>
active cooling. <o:p></o:p></p>
<p class="MsoNormal">If you have no thermal management solution, you do not have a viable design for
hypersonic flight!<o:p></o:p></p><p class="MsoNormal"><br /></p><p></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com8tag:blogger.com,1999:blog-2675974463524895416.post-73704248573434622742023-08-31T13:43:00.000-05:002023-08-31T13:43:27.998-05:00Famous Quote Is Still True<p>“Politicians and diapers must be changed often, and for the same reason” -- Mark Twain</p><p><br /></p><p class="MsoNormal"><o:p></o:p></p><p class="MsoNormal"><br /></p><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhXpUaZK8AVnHK1B4PpVIr1O_Kwo1l8v5LEOsljgQwtxQXzBf1iNOxlxbL1p4Sy5zyWEUeBQGyo8M30dRYoJlSquuJeJKiJHxXolQ9gcIZmq9iR9QiSeUgdCiA8_DnI05VRjR0eqd_X1iuupfS6l2i2GokazB-lYK9lj7MDzL03MXnTY6Lp8tVXDp95_B8V/s194/dirty%20diapers.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="130" data-original-width="194" height="268" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhXpUaZK8AVnHK1B4PpVIr1O_Kwo1l8v5LEOsljgQwtxQXzBf1iNOxlxbL1p4Sy5zyWEUeBQGyo8M30dRYoJlSquuJeJKiJHxXolQ9gcIZmq9iR9QiSeUgdCiA8_DnI05VRjR0eqd_X1iuupfS6l2i2GokazB-lYK9lj7MDzL03MXnTY6Lp8tVXDp95_B8V/w400-h268/dirty%20diapers.jpg" width="400" /></a></div><div><p class="MsoNormal">(Everybody knows what these are full of.)<o:p></o:p></p></div><div><br /></div><br /><div class="separator" style="clear: both; text-align: center;"><a href="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhhqxTF-YccLodZ14EPEzhld5YUVdPLWhgIofLTzgxhD_VAnOQ2ERwaYV7F1Ud_r9QUcbZmzSL8iqO77RfXzGaAklUCbzj6sAhiBhkrLiQmZ4DvIfMeoWUZf4IWEwxFhWGW73hMyujLN2or0hTFdVfWaSgKKN82ai9yK2hdDh0HSJOYRmIZB3bLGSU9k_ZC/s1024/Trump%20mug%20shot.jpg" imageanchor="1" style="margin-left: 1em; margin-right: 1em;"><img border="0" data-original-height="1024" data-original-width="1024" height="400" src="https://blogger.googleusercontent.com/img/b/R29vZ2xl/AVvXsEhhqxTF-YccLodZ14EPEzhld5YUVdPLWhgIofLTzgxhD_VAnOQ2ERwaYV7F1Ud_r9QUcbZmzSL8iqO77RfXzGaAklUCbzj6sAhiBhkrLiQmZ4DvIfMeoWUZf4IWEwxFhWGW73hMyujLN2or0hTFdVfWaSgKKN82ai9yK2hdDh0HSJOYRmIZB3bLGSU9k_ZC/w400-h400/Trump%20mug%20shot.jpg" width="400" /></a></div><p class="MsoNormal">(He is just one of very many, all full of the same thing as those diapers.)<o:p></o:p></p><p class="MsoNormal"><br /></p>Gary Johnsonhttp://www.blogger.com/profile/06723964751681093047noreply@blogger.com0