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.
I knew the hand-waving about combined-cycle turbine-scramjet
was BS, because about the fastest
practical speed for gas turbine is about Mach 3.2 to 3.3 due to overheat damage, and about the min takeover speed for scramjet
is Mach 4. Plus, the inlet and nozzle geometries are utterly
incompatible.
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, then operate
as a (subsonic-combustion) ramjet to accelerate above Mach 4, then finally operate as a scramjet to “fly
hypersonically” at or above Mach 5. The
ramjet and the gas turbine share similar inlet and nozzle geometries, but the scramjet is still utterly
incompatible geometrically with the other two. And,
you must change engine type in order to slow down for a more economical
cruise.
My suggested solution has, up to now, been “parallel-burn” propulsion: do not try to combine the various
propulsion types into one design,
instead install all 3 separately,
each optimized for what it is. (Combined, it is inevitable that performance of
each component suffers greatly.) But, a major problem with parallel burn at higher
speeds (where drag is high), is that no
one of these propulsive items is a large enough fraction of the vehicle frontal
cross section area! That severely limits
the max speed attainable, likely to less
than hypersonic, which eliminates any
reason to have the scramjet at all!
Concept for Combining Gas Turbine with Ramjet and
Scramjet
I have since had a sort-of hybrid idea. The 3 systems can share one common supersonic
inlet capture installation, but nothing
else! The post-capture channels of the
inlet must be made variable geometry, so
that the gas turbine and the ramjet can be fed subsonic air in a diverging
channel, while the scramjet is fed
supersonic air in a constant-area channel.
The supersonic channel to the scramjet must be “straight
through”, you absolutely cannot
divert a channel carrying supersonic flow,
because the turn always causes shock-down to subsonic flow! Anybody who claims otherwise is spouting pure
BS!
The gas turbine needs to be a low-bypass ratio afterburning
design suitable for supersonic flight,
and also be fitted with air bypass tubes around its core big enough so
that they can carry 100% of the air flow,
tapped off ahead of the compressor face, and going directly to the afterburner. (In the SR-71, those engines had 25% max air bypass, tapped from the 3rd or 4th
stage of the compressor.) In that way
with 100% bypass, the afterburner can
also serve as the subsonic-combustion ramjet combustor, using the very same post-capture subsonic
inlet air channel as the turbine uses. But, we do need to stop the airflow into
the compressor, to avoid overheat damage! And we need to stop backflow from the
afterburner into the turbine! Ramjet
combustor gas temperatures are far higher than any allowable turbine inlet
temperatures, and “leaks” lower the
ramjet pressure, lowering performance
drastically.
Therefore, it is a key
requirement here, when operating as a
ramjet, to stop the backflow from the
afterburner chamber from going up through the turbine into the turbine
engine. That is a serious and extremely difficult
design problem to solve! But it must
be solved, to prevent turbine
overheat, and to raise the achievable
chamber pressure of the ramjet, in order
to preserve its performance. Leaks are
low chamber pressure, and low pressure
is low performance. Period. That was settled long ago in tests.
What you “buy” with the 100% bypass and the backflow
stoppage complications, is a gas turbine
and a ramjet that share the same portion of the vehicle frontal cross
section, which then can be a much larger
fraction of vehicle frontal cross section,
so that the top speed in ramjet can be higher, reaching the scramjet takeover range at Mach
4+.
For scramjet takeover,
you must suddenly change the inlet post-capture channel geometry to a long, straight supersonic feed to the scramjet, that is also the “isolator duct” required for
stable scramjet operation. This scramjet
must be parallel-mounted to the rest of the propulsion, and must be completely separate, except for sharing the supersonic capture
features. It lets you put the scramjet
on the belly of the aircraft, and to use
the vehicle aft underside as a free-expansion nozzle surface. That reduces (but
does not zero) the scramjet’s fraction of the vehicle frontal cross section, as opposed to that of the turbine/ramjet, to about a 50-50 split. That highly-integrated geometry in turn increases
the max scramjet speed against drag,
making more-than-minimum (Mach 5) “hypersonic speed” feasible.
Doing these required design features is a
hellaciously-difficult problem, but does
offer a potentially-feasible solution for hypersonic flight that does not
involve rocket thrust to takeover speed.
I have not even touched on the thermal management issues, which may, in point of fact, be fatal to the concept! Suffice it to say the usual construction
techniques for the afterburner and its nozzle cannot be used, because for Mach 3.3+ speeds, there is no such thing as the cooling
air that those technologies require.
Finally, if the
marketing hype you see does not include a propulsion system that addresses the
issues I have raised here, and a
thermal management scheme that addresses the propulsion and the inlet and
the airframe, then I suggest that you
dismiss it as the BS that it quite evidently is!
A cartoon sketch of my scheme is given here as Figure 1.
Figure 1 – A Possible Means to Combine Gas Turbine Takeoff and
Landing with Scramjet Dash
Rocket-Boosted Ramjet Is a Much Better Way
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 much more economically 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 is hypersonic! No scramjet required! It just has lower specific impulse at
hypersonic speeds, as does the scramjet.
However,
you do not have to change propulsion to slow down to cruise!
If you include some small liquid rocket propulsion, your landing is not entirely “dead-stick”. 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!
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.
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 only
once per mission! The smaller liquid
rockets let you fly the plane at speeds below ramjet speed, for the approach and landing.
See Figure 2.
Figure 2 – Rocket-Boosted Ramjet as a Means to Achieve
Hypersonic Dash
Figure 3 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).
Combustor and nozzle heat protection is by ablative
materials, which cannot be re-used. So, the
IRR unit must be replaced 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.
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.
Figure 3 – Cartridge-Loaded Ramjet Combustor with IRR
Booster
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.
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.
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.
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!
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.
Related Information:
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.
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.
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.
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.
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.
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.
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.
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.
One Final Note:
All of this was done with open sources! 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.
--------------
Update 9-5-2023: I took some time to rough-out the characteristics of a rocket ramjet airplane design, 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, given in Figure 4. 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.
Figure 4 – Concept Flight Profiles
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 simplicity
above all else, 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.
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.
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 shown in
Figure 5, including the small-rocket
frontal thrust density value, based on
its exit area.
Figure 5 – Roughing Out a Small Liquid Rocket System
Emphasizing Simplicity Above All
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.
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.
The drag and ramjet thrust density and specific impulse data
are given in Figure 6.
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 shown in Figure 7.
Figure 6 – Rescaling Ramjet Performance From Some Old, Limited Data
Figure 7 – An Older Grain Design Used To Rescale IRR Booster
Performance
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, as indicated in Figure 8. This is less ramjet frontal thrust density
than originally desired, which is
what limited the max dash speed to Mach 5.
The data include a preliminary weight statement and some estimated
component lengths. Gross cruise range
exceeds 3000 nmi, at 85 kft.
Figure 8 – Results for the Flat-Bottomed Airframe With
Bifurcated Inlet
The radial distance from vehicle outer mold line to the case
or fuel tank OD is a critical variable,
as well. There must be some
such distance, to isolate thermally
the hot lateral skins from the vessels containing fuel or solid
propellant. That would include some
high-temperature mineral wool insulation.
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. But that is about
all I can realistically squeeze out of this design concept! The strakes containing the bifurcated inlet
and small liquid rocket equipment are just too large, driven by the required air inlet duct branch
sizes.
That trend illustrates the crucial role frontal thrust
density plays in high supersonic, low
hypersonic flight. There is no
getting around this, it is quite fundamental.
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.
I re-ran this alternate configuration, getting the results shown in Figure 9. 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.
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.
Figure 9 – Results for the Round Section Mid-Wing Airframe
with Inlet Through Fuel Tank
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.
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.)
Also bear in mind that the next most important
feasibility item is thermal management.
Those calculations have yet to be explored.
---------------
Update 9-18-2023: 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.
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 given in Figure 1. 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.
Figure 1 – Thermal Analysis of a “Typical” Lateral Skin
Panel
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.
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.
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.
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.
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.
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 plus 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.
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 illustrated
in Figure 2.
Figure 2 – “Typical” Thermal Equilibrium Considerations for
a Leading Edge Piece
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.
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. See Figure 3.
Figure 3 – Leading Edge Piece Results
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.
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.
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 will require
active cooling.
If you have no thermal management solution, you do not have a viable design for
hypersonic flight!
.jpg)
