Wednesday, June 1, 2022

About Hypersonic Vehicles

For hypersonic flight vehicles of any kind there are two fundamental problems that require solutions:  aeroheating and propulsion.  Of these,  the aeroheating is the more important!  If the design concept does not have a thermal management solution for the extreme aeroheating,  then regardless of any propulsion,  the design concept has no credibility at all!  See References 1 and 2 for more information.

The simplest propulsion solution is rocket.  This can take two distinctly-different forms:  (1) a hypersonic glider vehicle propelled by a large ballistic missile that is staged-off after doing its job,  and (2) a tactical-size hypersonic missile with its rocket propulsion on-board (which does not preclude adding a small staged-off booster rocket).  The range and top speed of the big ballistic missile-boosted concept is not inherently limited,  while the tactical-size vehicle with on-board rocket propulsion is very limited by the weight and volume constraints pertaining to whatever launches it.

However,  if you have a rocket on board the hypersonic vehicle,  you will have to protect it from the heat that conducts inward from the hot lateral skins (and the nose tip and any leading edges).  Same is true of any payload-related items.  These skin surfaces will have to operate in the 1000-1500 F (540-820 C) range in order to radiate enough heat away,  to balance the aeroheating input at stratospheric altitudes,  since the inward conduction simply must be interrupted to protect the rocket (and any other) internal components. 

The maximum recommended service temperature for titanium is 750-800 F (400-430 C).  It is very definitely NOT a high-temperature material,  despite what so many seem to think based on its use in the SR-71.  That vehicle was limited to speeds under Mach 3.3,  it was definitely NOT hypersonic!

Failing that re-radiation balance,  you simply will have to actively-cool those skins!  Why?  It is probably very infeasible to dump such large amounts of inward-conducted heat into the rocket propellant,  particularly if it is solid propellant,  and for several very compelling reasons.  See Figure 1 below. 

Hypersonic Airbreathing Propulsion

If one uses airbreathing propulsion to extend the range of a tactical-size hypersonic vehicle,  that inherently opens multiple further aeroheating problem issues,  that are simply not faced by a rocket vehicle or a glider.  You will (at the very least) have air inlet features and a combustor and nozzle to consider. 

The air capture cowl is aeroheated both inside and outside,  but can effectively radiate only from the outside,  and opportunities for conduction are geometrically absent,  so a higher equilibrium material temperature than a lateral skin is simply inevitable!  Internal inlet ducts obviously cannot radiate to the environment,  and must not conduct inward,  so they will require active cooling!  The combustor and nozzle will also require active cooling,  being either within the airframe unable to radiate,  or else actively aeroheated on the external surface if exposed.  Figure 1 below illustrates these items,  too.

Your choices for airbreathing propulsion are really quite limited:  ramjet,  scramjet,  and some sort of combined cycle propulsion (rocket/turbine,  ramjet/turbine,  or scramjet/turbine).  Turbine alone will simply not work at hypersonic speeds:  the fastest operational gas turbine engine was the short-life design in the Mig-25,  at Mach 3.5 maximum in the stratosphere!  At Mach 5 in the stratosphere,  the captured air temperature exceeds most turbine inlet temperature limits without burning any fuel at all!


Ramjet works quite well at Mach 3 to 4 in the stratosphere,  and can be readily designed to survive Mach 5 speeds with the modern technologies.  If the vehicle is low drag and the ramjet engine is ~100% of the vehicle frontal blockage area,  Mach 6 is demonstrably attainable,  maybe Mach 7.  However,  this is possible only with great difficulty solving the Mach 6-7 aeroheating problems,  especially those associated with the inlet capture,  internal inlet duct,  and inlet duct-mounted fuel injection hardware!  The only currently-viable technology solutions for heat protection at conditions like these are one-shot ablatives.  See References 3 and 4 for lots more information about subsonic-combustion ramjet.


Scramjet may well now be almost ready-to-apply technologically.  It has flown experimentally,  but not yet in vehicles with full aeroheat protections in place.  These tests were conducted at altitudes so high that the heat transfer coefficients were reduced by the low air density,  reducing the severity of the thermal management problem for the experimental designs (X-43A and X-51A,  plus an earlier Australian test).  At such altitudes,  airbreather frontal thrust densities are too low to provide any climb rate,  or any acceleration capability in level flight.  These are 100,000 to 130,000 foot (30-40 km).  

That high altitude effect has very serious implications for using airbreathing propulsion for flight-to-orbit,  since the airbreather (any airbreather!) will always have insufficient thrust to fly,  as the air thins further,  just because the ambient pressure is so low!  This is really why the X-30 project failed!  (Rocket actually has slightly higher thrust at altitude than at sea level,  but only if a conventional nozzle is used;  see Reference 5 for why that last statement is true.) 

Combined-Cycle Propulsion Issues

There are two fundamental problems with any (and all) combined-cycle engine designs that use gas turbine as one component.  Problem (1):  the inlet diffuser and nozzle geometries required by the gas turbine are fundamentally incompatible with scramjet (but not necessarily ramjet),  as illustrated in Figure 2 below.  Problem (2):  there must be zero airflow through the turbine component,  once the max safe speed for it (only about Mach 3 to 3.5) is surpassed.  The risk is overheating the turbomachinery.

Regarding problem (1),  gas turbines require low subsonic delivery speed at the compressor face,  which means the post-capture inlet is a divergent diffuser duct that is nearly all-subsonic.  Ramjet demands something similar,  although not geometrically identical.  (See Reference 6 for an explanation of those differences.)  Scramjet demands a nearly constant-area “isolator duct”,  that is all-supersonic to its outlet!  The very small divergence in that “isolator duct” merely offsets boundary layer thickening.  Variable-inlet-geometry hardware is well known to be both voluminous and heavy.  See again Figure 2. 

The turbine outlet speed from a gas turbine is also generally subsonic.  The nozzle must neck-down to a minimum throat area to reach sonic speed,  and may have a very modest supersonic expansion ratio.  Ramjet demands something very similar,  but usually somewhat larger.  Jet fighters use moving “turkey feathers” that are air-cooled to accomplish the throat and exit area variations needed.  Scramjet cannot have a neck-down to a min-area throat,  only a supersonic expansion!   And,  at high supersonic and hypersonic speeds,  there is simply no such thing as “cooling air”,  which means no variable geometry nozzle technological solutions exist in anything resembling a ready-to-apply form!  Again,  see Figure 2.

Regarding problem (2),  there must be designed-in some way to bypass all (ALL!) the inlet air around the gas turbine component,  directly to the ramjet or the scramjet component.  Otherwise,  the hot high-supersonic and hypersonic air will simply destroy the turbomachinery,  even if it is not turning!  This diversion geometry is hard enough to do subsonically for a ramjet component,  and pretty-much technologically impossible to do supersonically for a scramjet component,  due to the shock-down risk.   And,  any such variable geometry inlet hardware is going to be voluminous and heavy,  as already stated.

Rocket-based combined cycles avoid the turbomachinery gas temperature problems.  These are essentially variations on the old ejector ramjet,  and conceptually could transition to scramjet,  if the ramjet uses a thermal choke instead of a physical convergence to a minimum throat area.  However,  no ramjet vehicle ever flew with a thermal choke,  instead always with a physical nozzle!  More than half a century ago,  tests clearly showed that thermal chokes resulted in too low a combustor pressure to ever get any effective performance out of the ramjet.

The true state-of-the-art for these combined cycle approaches is only concept design with finite-element computer analyses.  Not much real testing has been done,  and those results were always less than expected.  The design analyses were (and are) usually made with computational fluid dynamics (CFD),  which is still notoriously subject to both the garbage-in/garbage-out (GIGO) law,  and serious problems recognizing fully-converged numerical solutions. 

Here’s the real problem with CFD models:  there’s a lot more going on inside any engine than just compressible fluid flow with this-or-that turbulence model.  The physics of combustion are usually inadequately modeled,  and the physics of flameholding are usually NOT modeled at all,  in most CFD codes.  Yet these effects really dominate the physics in the engine!  The “gold standard” is thus still real test data with real hardware,  and there is actually precious little of that with most of these concepts. 

That test data objection applies to both the turbine-based and the rocket-based combined-cycle concepts.  These are thus nowhere-near ready-to-fly,  generally speaking.   Which in turn is why you cannot go to a propulsion company,  and just buy one off-the-shelf!  These notions get proposed a lot for government R&D funded efforts,  but none have ever completed any actual development programs.

Effective Propulsion Solutions For Hypersonic Flight

The real solution to these propulsive geometry dilemmas probably has more to do with “parallel burn” of separate propulsion devices,  than with any sort of combined-cycle engine approach.  Another name for “parallel burn” is “mixed propulsion”,  which craft such as the Douglas “Skyrocket”,  the NF-104,  and the XF-91 had.  This took the form of a rocket engine and a gas turbine engine,  on-board separately.

One possible example could be a ramjet vehicle with a built-in rocket booster,  but one able to burn both engines simultaneously after launch and at very high altitudes,  and at landing.  As separate propulsion devices,  the geometry and performance of each component can be optimized.  Forced to share otherwise-incompatible geometries,  both components will inherently end up far from optimal. 

Hypersonic Airbreathing Propulsion for Orbital Ascent?

The real problem is that thrust of an airbreather (any airbreather!),  and the vehicle lift and drag,  are roughly proportional to ambient atmospheric pressure,  while vehicle weight is not!  At very high altitudes in the thin air,  there is not enough lift to oppose the normal weight component,  and not enough thrust-minus-drag net force available to overcome the axial weight component,  and so thus the vehicle to fails to fly steady-state,  much less climb and/or accelerate.  This is shown in Figure 3. 

This effect is the source of the “service ceiling” for an aircraft powered by an airbreathing engine (any type of airbreathing engine!).  This is usually specified to be the altitude at which the rate of climb (R/C) falls under about 200 feet per minute,  which is 3.33 ft/sec vertical velocity.  At Mach 5 (about 5000 ft/sec velocity),  that would be a climbing path angle near 0.03 degrees.  At only 500 ft/sec flight velocity (V),  that would be a climb path angle nearer 0.30 degrees.  Neither is very discernible above horizontal. And THAT is the point here:  effectively that is no rate of climb capability,  no matter how long you try.

Lower down,  where the air is thicker,  these forces become far more favorable,  but the aeroheating and drag problems to overcome are far worse.  This is because the heat transfer coefficients are roughly proportional to atmospheric density raised to a fractional power near 0.8,  not 1.  That assessment comes from the usual formulation of the Nusselt number correlations from Reynolds number:  constant x Reynolds number-to-the-0.8 x Prandtl number-to-the-1/3,  for turbulent flow.  See again Figure 3.   

That aeroheating effect and the drag are what drive the need to be at really high altitudes as the vehicle speed approaches orbital values.  Yet for thrust-relative-to-weight purposes,  the airbreather needs to be at very much lower altitudes!  That fundamental design requirements incompatibility is quite stark! 

You overcome it by using rocket propulsion,  not airbreathing propulsion,  at those very high altitudes in the really thin air.  Or at least use rocket propulsion simultaneously with your airbreather.  Yes,  the airbreather makes its thrust at high specific impulse,  but it just does not make very much thrust in that thin air!  The rocket does.  Which is why one should prefer the rocket,  when leaving the atmosphere!

For a two-stage launch system where the first stage is an airplane of some kind,  there are three important variables to consider at your intended staging condition.  In order of importance,  they are (1) highest possible speed,  (2) path angle at about 45 (or more) degrees above horizontal,  and (3) highest possible altitude.  Speed has the greatest beneficial effect,  altitude the least. 

Path angle is important so that the second stage may fly a non-lifting gravity-turn trajectory at minimum drag loss,  to orbit.  Pulling up at high speed is a large-radius turn requiring a lot of gees and incurring a lot of drag loss due to a very high lift requirement.   If your first stage can do its trajectory to arrive at both high speed and path angle,  that minimizes the second stage drag loss and impulse requirement! 

But,  at all but the very lowest altitudes,  this will require rocket propulsion as the entire propulsion system,  or at least in combination with the airbreather in parallel burn.   No airbreather of any kind used alone will be able to do this kind of beneficial first stage trajectory,  precisely because of the thin-air “service ceiling” effect at higher altitudes.

How Big A Threat Is This New Hypersonic Weapon Stuff,  Really?

So,  there are very good reasons why the new “hypersonic weapons” currently being ballyhooed in the press are just rocket-powered tactical missiles with a peak speed above Mach 5.  Otherwise,  they are just hypersonic gliders dispensed from a large ballistic missile flown on a low,  rather flat,  trajectory.  The “scramjet missiles” are still experimental items,  not really fieldable weapons,  for a while yet. 

The old,  retired AIM-54 “Phoenix” rocket-powered missile had a peak speed of just about Mach 5,  way back in the 1970’s.  So,  what is so “new” or threatening about reprising that?  Nothing!

We have had maneuverable re-entry vehicles as space capsules since the 1960’s,  and as ballistic missile warheads since the 1980’s.  The space shuttle was another.  There is nothing “new” there! 

For the “hypersonic gliders” to be much of a military threat,  the launcher has to fly a much lower trajectory,  at a very shallow angle below horizontal,  very unlike the usual strategic ballistic missile.  Otherwise, there is no time to maneuver the glider before it impacts.  That’s just high school physics!

This “new hypersonic glider threat” is really no big deal,  if you know to watch for those depressed ballistic missile trajectories.  And we do.  I just told you,  if no one else did.


In addition to the six references cited above,  there is a seventh very useful item,  for those who wish to research this topic further and deeper.  It is included in the list here as Reference 7.  That one contains lists of articles sorted by the topic area. 

One of those topic areas is “aerodynamics and heat transfer articles”,  where I put the high-speed aerothermodynamics stuff,  among some other things.  The hypersonics-related stuff is there,  right up to entry heating models.  There’s also topic areas for “ramjet” and “rocket” stuff,  and much more.  All of these are articles that I wrote and published on this “exrocketman” site over the last several years. 

To find any such article quickly,  use the navigation tool on the left side of the page.  You will need the posting date and the title (jot them down).  Click on the year,  then month,  then the title (if need be). 

You can click on any figure in an article to see enlargements of all of the figures in the article.  There is an X-out option at top right of that page,  which takes you right back to the article itself. 

#1. 2 January 2020,  “High Speed Aerodynamics and Heat Transfer”  (physics and calculation models)

#2. 12 June 2017,  “Shock Impingement Heating Is Very Dangerous” (physics with X-15 as an example)

#3. 10 December 2016,  “Primer on Ramjets” (basic concepts and fundamentals)

#4. 21 December 2012,  “Ramjet Cycle Analyses”  (how these things are best calculated)

#5. 12 November 2018,  “How Propulsion Nozzles Work” (covers conventional and free-expansion)

#6. 9 November 2020,  “Fundamentals of Inlets” (same components used quite differently for ramjets and gas turbines)

#7. 21 October 2021,  “Lists of Some Articles By Topic Area” (dates and titles arranged by topic)

Figure 1 – Heat Transfer Issues With Hypersonic Flight

Figure 2 – Geometric Incompatibilities Among Airbreathing Concepts For Hypersonic Flight


Figure 3 – Thin-Air Effects On Thrust,  Lift,  Drag,  and Aeroheating at High Altitudes

Addendum 6-11-22:

Here is a plot about the flight test space covered by the X-15 program,  relative to the aerodynamic ascent path to low Earth orbit.  The lower altitudes correspond to higher speeds reached by the X-15.  The higher altitudes correspond to lower speeds achieved.  

The X-15A-2 variant was able to reach 4520 mph (Mach 6.7) at 19.3 miles altitude,  still near the left end of that ascent corridor.  It was carrying a scramjet test article on its ventral fin stub during this flight.  The shock wave off the scramjet inlet compression spike nearly cut the tail off the bird from shock impingement heating effects. 

Now bear in mind that shock impingement effects multiply (considerably) the heating rates,  but not the plasma temperatures themselves.  All that means is that the structures affected are going to soak out very quickly,  to very near the plasma driving temperature,  pretty much regardless of what the designer might have done in the way of cooling provisions.  

At the X-15A-2 peak speed conditions,  that plasma sheath temperature was some 3070-3080 F,  far beyond the capability of even the Inconel-X skin material.  The white ceramic coating applied to this particular variant would have been able to do very little about this fast soak-out effect at the shockwave impingement locations.  So,  it is not surprising at all (in retrospect) that the craft suffered so much damage.  

What this figure says so very clearly is that if you fly too high,  you will not get enough lift to fly aerodynamically in such thin air.  And if you fly too low,  you will suffer unendurable aeroheating,  no matter how you attempt to construct your craft.  

And the X-15A-2 experience cited here says that you must avoid any shapes that might cause a shockwave from one structure to impinge upon an adjacent structure.  That absolutely rules out parallel-mounted nacelles.  

It is even worse near the orbital entry point end of the aerodynamic ascent corridor.  That would be about 17,000 mph at about 70 miles altitude.  The estimated plasma sheath driving temperature there is just about 13,200 F.  The peak heating point lies in between these two point (the X-15A-2 test and the orbital entry point).  

Your nosetip and leading edges will have to be ablatives.  And there is no allowing any significant soak-out for lateral skins.  Not at plasma temperatures in the 3000-13,000 F range!

By way of comparison,  the vertically-launched rocket vehicles we currently use to reach orbit (and that includes the Space Shuttle),  leave the sensible atmosphere at about 30 miles,  but only about 1500-2000 mph at that point.  They stay well above the aerodynamic ascent corridor,  essentially flying most of the way in vacuum,  without severe heating.

It is re-entry from orbit that essentially flies this aerodynamic ascent corridor in reverse.  It's just that the exposure times are a lot shorter for entry than they are for ascent.  So the overall heating exposures for aerodynamic ascent are much worse than for entry.

Those are the fundamental things you have to consider,  if you wish to fly an aerodynamic ascent to orbit,  regardless of the propulsion you may be considering. 

Update 6-12-22:

A lot of folks are using the word “hypersonic” very loosely.  Too loosely.  I have even seen the SR-71 termed “hypersonic”,  when it most definitely was not,  at a max allowable cruise speed of Mach 3.2 at some 85,000 feet.  Pilots would pull up and reduce throttle to slow down quickly,  if it ever reached Mach 3.3,  to avoid engine and airframe damage.  I have talked to a couple of them,  and that is what they told me.

There is a formal definition:  “hypersonic” is that speed at which the shape of the shock wave system about the vehicle no longer changes its shape appreciably as flight speed increases further.  The speed at which that happens depends upon the shape of the vehicle. 

For “pointy” shapes likes missiles and aircraft,  it is just about Mach 5,  or a velocity that is 5 times the speed of sound.  For really blunt shapes like space capsules traveling heat shield forward,  it is just about Mach 3.  Those two values are the rules of thumb for “hypersonic”,  for those two classes of shape.

Here is a further update to the figure showing the ascent corridor and the X-15 data.  I have added a typical vertical-launch trajectory,  and some effective temperatures in the plasma sheath about any vehicle.  I also show how precipitously the heat transfer coefficient drops as altitude extremizes.  That is why the peak heating rate point is somewhere in the middle of the ascent corridor,  not its high-speed end.