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
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
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.
References
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.
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.
