Sunday, October 27, 2013

Manned Launch to LEO Using Ramjet Missile Technology

Reported here is a concept for a two-stage to orbit (TSTO) vehicle,  capable of sending two astronauts to low Earth orbit (LEO) in a minimal transfer and return capsule.  Modern ablative heat shield items are assumed to reduce the capsule weight.  The capsule has a “service module” containing on-orbit propulsion and folding solar panels for electric power. 

These astronauts arrive in LEO with a significant on-orbit maneuvering propellant budget,  and can conduct short term (one week or less) missions unaided.  Longer missions would require docking with something that had some living space and life support supplies.  The on-orbit propellants are nitrogen tetroxide-monomethyl hydrazine (NTO-MMH),  which are very storable for very long periods in space.  Total on-orbit delta-vee was simply assumed to be 1 km/s. 

These roughed-out designs are the results of bounding calculations,  not trajectory simulations.  I went through four iterations to reach a rough size-out that I trust.  The second stage was analyzed with a loss-corrected approximation of the rocket equation.  The first stage was analyzed with rocket-impulse and ramjet-energy approximation analyses.  Everything was checked with thrust versus drag maps to ensure basic feasibility,  a considerable problem reaching high speed in the thin air where staging takes place. 

For the results of a true horizontal takeoff study utilizing ramjet assist,  see "HTO/HL Launch with Ramjet Assist",  dated 11-6-13.

First Stage

The first stage takes the form of two integral rocket-ramjet (IRR) booster pods,  based on a solid-propellant rocket booster,  and a liquid-fueled ramjet sustainer,  right out of missile technology based on ASALM-PTV,  about 1980.  The same technology is in the “Sunburn” and “Yakhonts” missiles built by the Russians today.  The booster is packaged within the ramjet combustor,  requiring some sort of inlet port cover,  and some sort of ejectable booster nozzle nested within the ramjet nozzle,  just as in ASALM and the Russian missiles.  10-30-13:  also the US ALVRJ (1970's),  AAAM (1990's),  and VDR (1980's-1990's) efforts used the same basic IRR technologies (I actually worked on ASALM,  AAAM,  and VFDR).  As did the 1960's Soviet SA-6 "Gainful" (that I also got to exploit ca. 1970's).  

I have previously posted some of the ramjet thrust performance data for such a pod (see references list),  based on RJ-5 fuel,  which is a synthetic similar to kerosene,  but slightly denser than water.   This work was based on a fixed ramjet nozzle geometry,  but a variable inlet geometry to maintain shock-on-lip throughout the range of ramjet flight speeds.  This amounts to a very simple translating compression spike,  similar to those used on the SR-71 “Black Bird”,  except operated for ramjet instead of turbine.  (The two engines are quite different,  even though the inlet components are the same.)

The IRR booster in each pod is a state-of-the-art AP-HTPB composite propellant of 20% aluminum and 87% solids.  It is similar to shuttle SRB propellant,  except for the really-large thixotropic effective-viscosity that is inherent with the really-high solids content that confers high Isp (some 254 demonstrated seconds).  Not many companies had the expertise to reliably cast such a propellant,  but the one I worked for decades ago did.  This is also essentially the same as the AIM-120 AMRAAM propellant we used successfully at the McGregor,  Texas plant,  before corporate politics closed it. 

Second Stage

The second stage is essentially Centaur technology,  revised to a slender form factor.  This is stainless steel balloon-tank technology,  with one engine similar to an RL-10.  It is a LOX-LH2 system.  I assumed only about 450 seconds of Isp,  in order to do a conservative performance estimate.  (Centaur actually does better than this.)  I assumed no need for wings during pull-up from staging,  relying entirely on body lift at Mach 6. 

Partial Reusability

I made the first stage booster pods reusable as flyback remote-control (R/C) aircraft,  fitted for a runway landing on land.  I didn’t get enough reserve fuel out of this bounding analysis to support flying back all the way to the launch site.  There are two of these pods,  one on each side,  which confers plenty of body lift,  even at only Mach 1.7,  the ramjet takeover point that I used.  The ramjet design works all the way down to Mach 1.5,  but I set takeover at Mach 1.7 to get some reliability “margin”.  After all,  this is supposed to be a manned system. 

The second stage is assumed expendable,  just as Centaur is today.  Its price ought to look about like that of Centaur.  The service module is also assumed expendable.  The capsule could be re-flown many times,  given about 2-to-3 inches (50-to-75 mm) thickness of PICA-X for its heat shield. 

First Stage Trajectory

This thing launches ballistically at a “high” gee level still tolerable by a human crew:  around 5 gees.  That’s fast enough to achieve flight speed “without time to swap ends”.  It comes off a slanted rail launcher (angle not determined here,  that’s for “real” trajectory code analysis),  and reaches ramjet takeover (Mach 1.7) at the relatively low altitude of about 5000 ft (1.5 km) above mean sea level (MSL).  Control during this phase would be by relatively-small attitude thrusters.  See Figure 1.

Once the solid integral boosters burn out,  the ramjets “light up”.  Transition in ASALM was 100 milliseconds,  this should be comparable to that.  Each pod has a frangible-glass port cover protected with bonded rubber insulation,  and equipped with a destruct charge set off as booster chamber pressure tails off.  Each pod has an ejectable booster nozzle (I computed roughly a 10 inch diameter throat,  a 38 inch diameter exit,  and a 52 inch exit bell length at 15-degree half angle).   All of this is based on ASALM technology.

I guessed these ejectable nozzle assemblies at 200 lb per pod,  scaled up from 15 lb for ASALM,  more or less on diameter squared,  not cubed,  which may or may not be as realistic as it should be.  These nozzles are held in place by a circular snap ring in a groove near the exit,  just like ASALM.  That snap ring is underlain with very-high-quality detonation cord,  also set off by the tailoff of booster pressure,  just like ASALM.  Then the fast-declining residual boost pressures “shove” the ejectable nozzle clear.  The boost pressure tailoff also sets off a magnesium flare to light the fuel-air reaction,  same as ASALM.  (From about Mach 3 on up,  the fuel-air reaction is essentially hypergolic.) 

The ramjets have very significant thrust margin over drag,  although vehicle accelerations in ramjet are far less (factor 10) than in IRR rocket boost.  The trajectory assumes a constant-altitude acceleration to Mach 2 at 5000 feet,  where thrust margin is pretty much optimum.  The vehicle then pulls up on body lift for a Mach 2 climb to 60,000 feet altitude (18.3 km).  I did a spreadsheet thrust-and-drag point-performance-estimate series,  to ensure this would be feasible.  These estimates used the same thrust and overall-efficiency results reported for the pod,  just correlated versus Mach number and altitude. 

Once the vehicle reaches 60,000 feet at Mach 2,  it pulls-over horizontal and uses its thrust margin to accelerate to Mach 6 for staging.  The crucial controlling factor for acceleration to this speed is vehicle drag.  It sets the diameter ratio of the second stage to each of the two IRR booster pods.  That is where the 7 foot to 5.22 foot diameter ratio comes from.  The ramjet cycle analysis that I used is beginning to break down from ionization/recombination effects,  from about Mach 5 on up.  It begins diverging sharply from reality above about Mach 6.  That and the ASALM speed record,  are where the staging speed used here came from.  Overall energy conversion efficiencies fall in the 22-23% class. 

This relative diameter result is a function of the number of pods.  That determines how much drag must be overcome by the ramjet thrust available in the “thin air” at altitude.    I got 5.22 ft diameter for each of two pods,  carrying a minimum-credible 7 ft diameter payload.  My drag data came from that for clean projectiles in the old Hoerner “drag bible”.  (If you don’t know what that reference is and how to use it,  you have no business playing with this concept.)

The pods have to be nose inlet configuration,  in order to limit pod drag to skin friction and a little cowl lip pressure drag (that is the minimum achievable).  Side inlet configurations will be “draggier”,  and will have far more difficulty reaching Mach 6 at 60,000 feet (or any other altitude).  The single nose inlet selection here is therefore the “optimum” configuration choice.  It features a cowl area about 50% of the pod cross section area,  an inlet duct about 40%,  a ramjet throat about 65%,  and a ramjet nozzle exit pretty close to 100% of the pod cross section. 

Higher staging altitudes reduce frontal ramjet thrust,  reducing vehicle acceleration,  extending the acceleration range,  and increasing the first stage fuel and article weights.  This is based on earlier trade studies I did with similar bounding calculations.  It’s a judgment call,  but 60,000 ft seems to be a pretty good “rule of thumb”.

Second Stage Trajectory

Since the staging point is Mach 6 at 60,000 ft,  there is plenty of body lift available to pull up to about 40-degree path angle.  “Store separation” hypersonically is even more problematical than at supersonic speeds,  something the military avoids at all costs.  That is because the aerodynamic forces are so much larger than the weight forces,  even in the thin air.  That means stage separation will require small solid motors strategically placed,  on both stages.  Those details are beyond scope here,  although the solid motors would resemble scaled-up versions of the ullage motors seen in many launch vehicles. 

The second stage pulls up to a steep path angle on body lift,  then flies a simple ballistic gravity-drag turn to orbit.  See Figure 2.  I simply assumed a 10% gravity-drag combined loss,  imposed on the velocity requirement for simple rocket equation mass ratio analysis.  Mass ratio sets the propellant mass fraction.  That and the inert fraction set the payload fraction. 

If the mass of the payload is fixed (as it is here),  that sets the weight statement,  for any given propellant combination.  The combination here is liquid oxygen-liquid hydrogen (LOX-LH2),  similar to that in the Centaur upper stage in use today.  As in Centaur,  I assumed steel balloon tankage construction,  just of a slim form factor.  Inerts are 10% of stage ignition weight,  as in Centaur. 

The Ultimate Payload

What is delivered to LEO is a manned capsule with a “service module” that supplies solar electric power plus significant on-orbit maneuver capability.  I assumed (1) very storable propellants,  and (2) 1 km/s maneuver capability.  The propellants are NTO-MMH.  I assumed a conservative Isp of 300 s for the maneuver engine.  See Figure 3. 

The capsule is a minimum-credible two-man re-entry vehicle of an assumed 2000 lb,  assuming low-density ablative heat shielding that can be re-flown a few times.  That would be PICA-X for the main heat shield,  and Avcoat on the afterbody external surfaces.  The minimum credible diameter for a cramped two-person capsule would be near 7 feet. 

This capsule would contain nothing more than two persons,  their spacesuits,  and many hours of life support.  The idea would be adequate support for a few days of independent operation,  or a “lifetime” of many weeks to several months if docked to an independent habitat such as the International Space Station.  There are an entire plethora of useful missions that such a crew transport capability could support,  a subject not explored here.

Second Stage

The second stage is a simple rocket vehicle using LOX-LH2 propulsion,  with a gimballed engine.  I used a simplified analysis and a conservative Isp to estimate the stage properties shown in Figure 4.  The payload is the two-man capsule and service module already described as the “ultimate payload”.  This second stage is not assumed to be recovered and reused,  although there is always that possibility,  given the higher inert fraction required to cover the heat shield and parachute equipment,  and perhaps some additional deorbit/deceleration propellants.  Making that stage recoverable and reusable is far out of scope here.

First Stage IRR Pods

The first stage IRR pods are shown in Figure 5.  This pod concept is basically identical to that described in the earlier postings regarding ramjet pod performance.  The booster is AP-HTPB composite propellant under web fraction and volumetric-loading conditions easy to meet,  with a burn rate already well-demonstrated without resort to ultra-fine AP grinds.  The ramjet uses the dense synthetic kerosene-like fuel RJ-5,  sometimes known as Shelldyne-H. 

This is the nose-inlet analog to the chin-inlet ASALM-PTV missile configuration.  The only difference between these results and the earlier postings of ramjet pod performance is in the pod drag:  fins that fold into a wake zone,  instead of fixed fins.  Those fins are needed for stability and control,  after staging,  for recovery purposes.  They have significant drag,  left in the slipstream.

Recovery is by runway landing,  on a steerable nosewheel and fixed skids on the ventral fins.  Two pivoting wings are extended subsonically,  to form a biplane configuration capable of landing at a bit over 150 knots equivalent air speed (KEAS),  without resort to flaps and leading-edge devices.  These pods are not manned;  recovery is by remote radio control. 

They cruise back as far as fuel allows,  at Mach 2 60,000 ft conditions,  fins extended,  but wings stowed.  Once the engine is “off”,  the pod quickly decelerates subsonic,  at which point the wings pivot to extend for subsonic glide to a dead-stick landing on the runway. 

These pods are modular:  an inlet section,  a tank section,  and a combustor section.  Refurbishment of the pod requires separation of the combustor IRR section,  and its replacement by another already loaded with propellant and transition gear out of inventory. 

The inlet section merely requires minimal servicing of the hydraulics for the nosewheel and translating inlet compression spike.  The tank section merely needs fuel control checkout,  and tank refueling.  All three sections need the external cork layer residues power-washed off,  and replaced with fresh cork insulation.  This was a technique used very successfully on the Phoenix missile for hypersonic aeroheat protection.  

The combustor section needs the internal insulation residues power-washed clean.  Then it needs a new DC 93-104 insulator cast in place on retaining ribbons,  and primed for an etched Teflon separator sheet.  That sheeted insulator then needs another primer before casting the booster propellant.  This system was quite successful in ASALM,  in spite of the otherwise-incompatible chemistries of propellant and insulator.

The booster is envisioned as a simple internal-burning cylinder with both ends unrestricted,  which would be slightly progressive,  more-or-less matching the drag increase with speed.  The propellant resembles shuttle SRB propellant,  except for the higher solids content that makes it very thixotropic.  This material has to be pressure-cast into vacuum in the case,  which is fitted with cast tooling.  The pressure-casting pressure level is very significant,  or else voids will cause fatal troubles.  This was a processing technique well-proven at the old McGregor, Texas weapons plant.  Few others could do this.

The unique concept here is “circumferential folding” of the fins,  something fairly compatible with the limited-volume of the ramjet nozzle recess,  by means of circumferentially-rotating rings to which the fins are fixed.  During rocket boost and ramjet sustain,  the three fins of each pod are stowed in the wake of the second stage,  for no perceptible drag increment.  That is how power-on pod drag can be limited to skin friction plus a bit of cowl lip pressure drag.  (Power-off,  one must include the base drag,  too,  which differs between rocket and ramjet,  due to the differing exit area sizes.)    

Upon staging,  the fins are quickly moved to their proper positions.  This can be done hypersonically,  because each fin is essentially streamline to the flow,  and thus is not subjected to catastrophic forces.  The pod flies supersonically on body lift until fuel is exhausted.  Deceleration to subsonic is then quite rapid,  after which the pivoting wings are extended.  Two wings are used,  in a biplane configuration,  to reduce the wing loading under 100 lb/sq.ft,  which gets the sea level stall speed down near 150 knots equivalent air speed.  That is a very practical value,  without any flaps or leading-edge devices.  Each wing is about 35 foot span,  and constant 2 ft chord.  A flat-bottomed subsonic airfoil is the proper choice. 

Cluster Vehicle

The vehicle comprises the second stage rocket plus two IRR pods strapped-on,  as shown in Figure 6.  The strap-ons are positioned on either side for maximum body lift capability.  They are staggered axially to provide the room at the rear in which to stow the fins in the wake of the second stage.  The wings are stowed top and bottom by pivoting into axial alignment with the pod airframe. 

This is necessarily a cluster vehicle.  The axial stagger ensures that the second stage bow shockwave does not impinge upon the strap-on IRR pods.  However,  the spike shocks of the pod inlets do impinge upon the second stage.  Depending upon material selections,  the enhanced shock-impingement heating can be quite catastrophic above about Mach 2-to-3.  Even with heavy Inconel-X skins,  shock-impingement heating became quite catastrophic on the X-15 flight that carried a scramjet test device in place of its ventral fin. 

For this vehicle,  small sacrificial panels of the tough ablative silica phenolic could be located in the impingement zones on the second stage.  Although expensive and a bit heavy,  such panels would provide a reasonably-certain “fix” for the otherwise-fatal shock-impingement heating problem. 

Concluding Remarks

I did not even try to estimate costs.  But,  consider this:  this vehicle can send a crew of 2 to LEO,  yet grosses near only 120,000 lb at launch.  That is a lot smaller than current commercial satellite launch vehicles,  and more in line with the size of the historic Titans that launched Gemini in the 1960’s. 

The second stage rocket is a form-factor variant of the venerable Centaur upper stage,  and thus should be priced similarly. 

The small capsule and service module should resemble the historic Gemini of the 1960’s except that this one is partially reusable.  The price should be similar to Gemini corrected for inflation,  but significantly reduced for experience gained since then,  and for reusability of the return capsule itself. 

The first stage pods resemble nothing so much as large tactical missiles,  and should have similarly-small manufacturing and logistical support “tails”.  The prices for these should resemble the prices for large tactical missiles,  quite a bit cheaper than is “customary”,  even with commercial launchers. 

There is absolutely nothing here that requires new technology development.  There is only the application of well-established missile technology in a launch venue not accustomed to using it since about 1960. 

Just as a wild guess,  call it $25M per launch,  for 3300 lb delivered to LEO.  That’s about $7500/lb or $16,700/kg,  which is higher than commercial,  but still less than the trend of government-operated vehicles.  The amazing schedule flexibility of a missile-like launcher is preserved in this design,  a unique advantage indeed. 

My point is that simplicity,  flexibility,  and smaller launched size,  can all be had for a price not all that out-of-line with modern commercial launch. 


Sighard F. Hoerner,  “Fluid Dynamic Drag”,  self-published by the author and then his widow Liselotte,  1965.
Pratt and Whitney,  “Aeronautical Vest Pocket Handbook”,  12th edition,  December 1969 (propellant and performance data for liquid rockets).

G. W. Johnson,  “Inlet Data for Ramjet Strap-On Pod”,  posted to,  2-20-2010.

G. W. Johnson,  “Ramjet Strap-On Pod Concept”,  posted to,  2-20-2010.

G. W. Johnson,  “Ramjet Strap-On Pod Point Performance Mapping”,  posted to,  2-20-2010.

G. W. Johnson,  “Preliminary Acceleration Margins for Baseline Pod”,  posted to,  2-28-2010.

G. W. Johnson,  “More Ramjet Performance Numbers for the Strap-On Pod”,  posted to,  7-11-2010.

G. W. Johnson,  “More Strap-On Pod Ramjet Engine Data”,  posted to,  7-23-2010.

G. W. Johnson,  “Two Ramjet Aircraft Booster Studies”,  posted to,  8-22-2010.

Update 10-30-13:

The technology described here is one of several possibilities for flexible,  rapid,  and modestly-priced launch of a small work crew to orbit,  to do a specific job and then return.  The point of this article was to show that this goal can be achieved with technologies we have had for decadesAll that is needed is the will to do it,  not endless gravy-train technology-development programs,  nor politically-dictated gigantic rockets. 

The technology outlined in “End of an Era Need Not Be End of a Capability”,  posted 8-2-11,  is essentially the deployment on-orbit,  of generalized repair and maintenance facilities.  These are only occasionally manned,  when there is a job to do.  A few of these facilities would need to be deployed in a few convenient orbits. 

You send up propellants and life support supplies in small payloads when you want to use one of the repair facilities.  The propellants and some of the supplies might even be launched at high gee with a light gas gun,  an even-cheaper-to-use technology.    The crew collects these,  and brings it all to the repair station,  when they go up.  That’s why I gave the capsule a service module with a significant on-orbit maneuver capability.

If you combine a technology like this article’s crew transport capability,   with a permanently-deployed but occasionally-manned repair craft like that described in the 8-2-11 article,  then you have recreated the tremendously-successful and valuable on-orbit service and repair capability we had with the space shuttle.  

The changeand it is a huge one,  is that you no longer incur the gigantic expense of a shuttle launch (or any other giant rocket),  every single time you need to use the on-orbit repair capability. 

This approach is not the way our government space program has historically operated.  It is more like what the commercial space companies are trying to do,  but have not really accomplished yet.  Maybe there is a lesson for government space agencies heresince they still set many of the missions to be done,  and the rules by which to accomplish them.  

Use the wrong rules or goals,  and you get a bad result;  it’s just plain old common sense.  

Figure 1 – First Stage Trajectory Characteristics

Figure 2 – Second Stage Trajectory Characteristics

Figure 3 – The Ultimate Payload (Carried By the Second Stage)

Figure 4 – Second Stage Characteristics

Figure 5 – IRR Pod Characteristics

Figure 6 – Launched Vehicle (and Recovery) Characteristics


  1. Thank you for this post. i understand some topics and some of them are not but i am still trying to learn those point. Thank you.

    1. I'm glad you liked this one. I worked hard on the calculations. -- GW