Basic Thermal Results for High Speeds

This article is a direct follow-on to the updates posted to “Purported SR-72 Propulsion”,  posted 1 September 2023.  As I have said there,  and multiple places and times elsewhere,  if you do not have a thermal management design concept,  you do not have a feasible hypersonic flight concept!  This article attempts to put some bounds on that problem.

Lateral Skin Study

The following is a simplified equilibrium skin panel surface temperature estimate for lateral-facing skin panels.  These could be on aerosurfaces (wings and fins),  or on the sides of a fuselage body.  I did not consider any conduction inward or to adjacent panels.  I did not consider any active cooling.  There is convection to the panel,  and thermal re-radiation from it.  It soaks out hot enough to balance the two. 

I did this for Mach numbers from subsonic to Mach 7,  using standard compressible flow methods and the high-speed heat transfer models that are based upon it.  I used free-stream conditions as the good approximation that they really are,  for local edge-of-boundary layer conditions.  I did not analyze past Mach 7,  because the fundamental assumptions underlying compressible flow analysis methods are breaking down,  due to ionization into something that is no longer air as we know it.

I show temperature curves in Figure 1 for air total temperature,  boundary layer recovery temperature (the driver for heat transfer to the panel),  and equilibrium panel soak temperatures for low and high thermal emissivity.  The service temperature limits for a variety of materials are also shown.  Figure 2 shows the film coefficient trends vs Mach at 40 kft,  for low and high emissivity.  Beyond about Mach 3 or 4,  these are pretty constant.  Data in the same formats for 85 kft are in Figure 3 and 4,  and for 130 kft Figures 5 and 6. 

Figure 1 – Skin Panel Soak-Out vs Mach at 40 kft

Figure 2 – Film Coefficients vs Mach at 40 kft

Figure 3 – Skin Panel Soak-Out vs Mach at 85 kft

Figure 4 – Film Coefficients vs Mach at 85 kft

Figure 5 – Skin Panel Soak-Out vs Mach at 130 kft

Figure 6 – Film Coefficients vs Mach at 130 kft

Skin Study Correlation:

Recovery temperatures do not change so drastically with altitude,  unlike film coefficients.  See Figure 7.

Figure 7 – Replots of Film Coefficient and Soakout vs Altitude at Mach 5

As the figure shows,  the result is a drastic change in soakout temperatures,  driven by drastically lower film coefficients at extreme altitudes.  The recovery temperatures all fall between 3800 and 4500 F at Mach 7,  as shown in Figures 1, 3,  and 5 above.  This suggests that a single analysis could establish a representative film coefficient value insensitive to changes in speed,  at Mach 4+ and some altitude,  which could be quickly scaled to other altitudes.  Calculating recovery temperatures at each flight condition is a far easier thing to do.  The correlation supporting that shortcut is given in Figure 8.  Doing it that way is only a ballpark estimate that supports better,  more detailed analyses later.  But it is useful. 

Figure 8 – Correlating High-Speed Film Coefficient vs Altitude

Leading Edge Stagnation Study

There is a compressible flow-based heat transfer correlation for stagnation zone heating.  It exists in two forms,  determined by a coefficient on the Nusselt number expression:  C = 1.28 for nose tips,  and C = 0.95 for aerosurface leading edges.  I looked at leading edges for this study,  so bear in mind that nose tips will run a little hotter still.   

In this Nusselt correlation,  you evaluate boundary layer properties at the total pressure and total temperature properties behind a normal shock at flight conditions.  I used the NACA 1135 tables for this.  It also uses a second viscosity evaluated at the flight conditions.  I did this for Mach 2 to Mach 7,  at the same three altitudes as the skin panel study.  The idea was to balance convective heating against thermal re-radiation,  with no conduction or active cooling,  as in the skin panel study. 

The results at 40 kft are given in Figures 9 and 10.  Figure 9 shows trends of total temperature,  and two local stagnation-region equilibrium temperatures,  one at low emissivity,  one at high emissivity.  Figure 10 superposes material service limits on the same curves.  The same data in the same format is given in Figures 11 and 12 at 85 kft,  and Figures 13 and 14  at 130 kft. 

Figure 9 – Stagnation Region Soakout Results vs Mach at 40 kft

Figure 10 – Soakout at 40 kft with Service Limits,  and a Speed Limit Indicated with Inconel X-750

Figure 11 – Stagnation Region Soakout Results vs Mach at 85 kft

Figure 12 – Soakout at 85 kft with Service Limits,  and a Speed Limit Indicated with Inconel X-750

Figure 13 – Stagnation Region Soakout Results vs Mach at 130 kft

Figure 14 – Soakout at 130 kft with Service Limits,  and a Speed Limit Indicated with Inconel X-750

In Figures 10,  12,  and 14,  I have included data for the service temperature limits and tensile strength at those limits,  as part of the figure.  Of the metals possibly useful for these high speed exposures,  Inconel X-750 is by far the strongest,  leading to thinner parts of lower weight.  So,  I used it as the selection here,  for “best” performance.  Under the earlier name “Inconel-X”,  this was in fact the skin material and leading edge for the X-15 rocket plane,  which skin was a major load-bearing portion of its airframe. 

Even so,  the speed limit for Inconel X-750 in a stagnation zone is only about Mach 4.9 at 40 kft,  about Mach 5.2 at 85 kft,  and about Mach 5.8 at 130 kft.  For lateral skins,  this was nearer Mach 6 at 40 kft,  Mach 7 at 85 kft,  and likely near or above Mach 8 at 130 kft,  because the convective heat to be reradiated is far lower for lateral skins,  compared to stagnation zones. 

A good guess says the stagnation limit for Inconel X-750 is about Mach 5.5 at 100 kft,  which neatly explains why the X-15A-2 with the drop tanks was coated all-over with an ablative for its flights to Mach 6 and beyond,  despite the indicated survivability of its lateral skins at Mach 7+,  near 100 kft.  

The craft reached Mach 6.7 at 99,000 feet on flight 188,  and suffered shock-impingement heating damage to the underside of its tail,  to both lateral and stagnation surfaces.  That phenomenon drastically raises the local heating rate,  but not the actual gas temperatures,  as described in another of my articles on this site:  “Shock Impingement Heating Is Very Dangerous”,  posted 12 June 2017.  See also NASA TM-X-1669 ““Flight Experience With Shock Impingement and Interference Heating on the X-15-2 Research Airplane”,  dated October 1968,  and written by Joe D. Watts,  at the Flight Research Center,  Edwards,  CA.  This document is publicly available over the internet.

Stagnation Study Results:

Use no metals for leading edge stagnation zones that are cooled only by re-radiation,  past about Mach 5.5,  and then only above 100 kft.  You must instead use ablatives,  or apply massive active cooling.  See Figure 15.  

Figure 15 – Results for Stagnation Zone Equilibrium

Nose tips will run slightly hotter than leading edges (higher h values at the higher C raise Tsurf),  thereby have a somewhat lower speed limitation than leading edges.  The risk with both locations is distortion and collapse of the parts,  as they weaken rapidly with increasing overheat. 

Alloys like Rene 41 and Alloy 188 can take slightly higher temperatures than Inconel X-750,  but are inherently weaker structurally by around a factor of 2.  This is a crucial consideration,  because stagnation zones see the highest positive surface pressures on the airframe.  Distorted or failed leading edges lead to higher drag,  loss of lift,  and intrusion of hot gas inside the aerosurface,  something to be assiduously avoided.  In general,  weaker is thicker,  which is heavier. 

Lateral Skin Results:

Speed limits versus altitude for Inconel X-750 lateral skins are about Mach 6 at 40 kft,  a bit over Mach 7 at 85 kft,  and likely above Mach 8 at 130 kft.  This is complicated by the risks of shock impingement heating,  which occurrence is complex and difficult to predict,  and which can do fatal damage at much lower speeds nearer only Mach 6.  See Figure 16.  Bear in mind that the analysis method is invalid above about Mach 7,  although the prediction is likely still crudely true. 

Figure 16 – Results for Lateral Skin Equilibrium

As with stagnation zones,  there are alloys that will go a little hotter,  but at far lower strength.  This is a crucial consideration,  because in monocoque construction,  the skins are an integral part of the airframe structure,  bearing much more than just local surface pressure loads.  Weaker is thicker,  which is heavier.

Remarks About Airbreathers:

Components associated with airbreathers (of any type) were not studied here.  The X-15 was a rocket plane.  The results above apply to both rocket-powered hypersonic vehicles,  and to hypersonic gliders. 

All airbreathers will have some sort of supersonic inlet capture structures,  some sort of post-capture air ducting that leads to the engine device (whatever it is),  and that engine device and its nozzle.  The ducting,  engine device,  and nozzle might be either buried inside the airframe,  or exposed as part of the airframe.

               Air Inlet Components

Inlet capture features suffer worse heating effects than leading edge (or nose tip) stagnation surfaces,  This is because they are heated (unequally) on both outside and inside surfaces,  but can re-radiate to cool from only the exterior surfaces,  with very localized stagnation soak-out on leading edges that must stay thin and sharp,  in order to function properly.  There is little opportunity for any conduction-as-cooling,  and not much opportunity for any active cooling.  They must also contain serious internal pressures without shape distortion. 

Buried subsonic inlet ducts will inevitably soak out to essentially the full air recovery temperature,  or else  they must be actively cooled.  They cannot re-radiate,  being buried inside the airframe.  They must be externally insulated to protect the rest of the airframe and its contents. 

Exposed inlet ducts are unlikely in hypersonic designs,  as too much airframe drag gets added.  However,  these are also internally heated,  and can only re-radiate to cool from that portion of the outside surfaces not inside a fairing or facing the fuselage.  They will still tend to approach air recovery temperature soak-out,  although not as closely as buried ducts. 

               Combustor and Nozzle Components

Buried or exposed combustors eventually soak out to something in between the external and internal recovery temperatures,  and will likely need active cooling.  The buried combustor will take a longer time to equilibriate,  because it starts off exposed to low airframe internal temperatures,  with a relatively low thermal conductivity for the free convection or insulated interfaces between it and the skin.  But it will soak out very hot! 

An exposed combustor can re-radiate directly to the surroundings,  while the buried combustor cannot (while the airframe skin can),  so the exposed combustor may possibly equilibriate a little cooler than the buried combustor.  But neither has a cold “sink” to dump heat into.  They both get very hot! 

The same applies to propulsion nozzle structures,  whether buried or not. 

               Turbomachinery

As for turbomachinery (compressors and turbines),  these must be isolated completely from hot intake airflow above about Mach 3 to 3.5.  Beyond that speed,  the very intake air temperature exceeds the turbine inlet temperature limits of almost any conceivable design.  The main flying examples of these speed limitations were the XB-70 (Mach 3.0),  the SR-71 (Mach 3.2),  and the Mig-25 (Mach 3.5).

               (Subsonic-Combustion) Ramjet

Ramjet can fly faster than turbine,  before hitting overheat speed limits.  Flight tested but not fielded as operational,  the ASALM-PTV test vehicle was designed to cruise steady state at Mach 4 and 80 kft,  followed by an average Mach 5 terminal dive onto its target.  It did so successfully in flight test. 

In one test of ASALM-PTV,  an assembly error led to a throttle runaway incident,  with the vehicle accelerating to fuel exhaustion at Mach 6 at low altitude (near 20 kft).  It suffered airframe overheat damage,  but actually survived the short transient flight and was recovered after it crashed. 

If designed for it,  ramjet could conceivably be made to work steady-state at Mach 6,  or even a bit faster,  perhaps.  The internal air duct and combustor/nozzle will require active cooling for a long flight.  The inlet cowl lip surfaces will likely need to be made of a really high-melting metal,  like tungsten or columbium,  so that they remain both sharp and thin,  without distorting.

               Supersonic-Combustion Ramjet (Scramjet)

Scramjet can fly faster still than ramjet,  but faces similar overheat risks for its inlet capture and supersonic isolator duct,  and its combustor and nozzle structures.  These get ridiculously difficult to design for,  as speeds increase beyond Mach 7.  The same can be said for airframe stagnation surfaces and lateral skins.  Short transients and ablative materials make such flight possible,  but those are neither reusable,  nor are they long-range. 

               Altitude Limits

The problem with all airbreathers,  of any type whatsoever,  is the “service ceiling” effect.  These devices produce an altitude-dependent characteristic trend of thrust versus speed,  with lower thrust levels in the thinner air at higher altitudes.  Roughly speaking,  thrust is proportional to the ambient atmospheric pressure at altitude.  So is drag.  But weight does not vary with altitude,  only with time as fuel burns off.

The vehicle requires enough lift to offset the perpendicular component of its weight,  as it tries to fly up an ascending path.  It also requires enough thrust to offset the sum of drag and the pathwise component of its weight.  See Figure 17.

Figure 17 – Why There Is an Altitude Limit for Airbreathers

There is an altitude at which there is insufficient thrust to overcome drag and the weight component,  regardless of any wings that might solve the lift problem.  Above that altitude,  it cannot even fly level steady-state,  at all.  As a rule-of-thumb at speeds in the Mach 5 to 7 range,  that’s around 130 kft,  almost no matter what sort of airbreather you might design.

Remarks on Active Cooling

This can be done reusably with a dedicated liquid coolant,  or it can be done regeneratively with the fuel.  For rocket systems,  the oxidizers are not generally very good coolant materials,  while the fuels generally are.  Either way,  the coolant may not be allowed to boil inside the cooling passages,  because that leads to vapor lock and a stoppage of coolant flow.  That in turn requires you to operate your coolant passages at very high pressures to avoid boiling,  which costs weight,  and power to run. 

However,  even if you deliberately allow boiling,  that reduces heat transfer capacity of the coolant,  because the gas density is so much lower than the liquid density,  for all known coolant materials.  This is really a per unit volume problem,  rather than a per unit mass problem,  because the passage sizes are pretty much fixed. 

Final Remarks

What I have done here is bound the problem for rocket-propelled vehicles,  or gliders,  that fly hypersonically.  I did this in terms of steady-state equilibrium surface temperatures,  for lateral skins,  and for stagnation zones on nose tips and aerosurface leading edges. 

I have provided some discussions,  but no numbers,  for the airbreathing propulsion components that might be applied to hypersonic vehicles.  Those are worse to thermally-manage than stagnation zones.

I have commented upon the “service ceiling” effect that applies to any airbreather of any kind at all.  This is related to the narrow flight corridor to orbit,  that resulted from the X-15 program.  See also “About Hypersonic Vehicles”,  posted 1 June 2022,  on this site.  Plots of that corridor are in that article.

And I have commented upon the difficulties faced by any actively-cooled designs.  

Note:

This article has been included in the catalog article,  under the topics “aerothermo” and “ramjet”.  That article is “Lists of Some Articles by Topic Area”,  posted 21 October 2021.  The fastest way to reach it is to use the navigation tool on the left side of this page.  To use it,  you need the article posting date,  and its title,  so in general,  jot that stuff down.  Click on the year,  then on the month,  then on the title if more than one item was posted that month.  Simple as that.