The document discusses the advancements and challenges of hypersonic travel, particularly focusing on air-breathing propulsion systems like scramjets. It highlights the technological capabilities, flight history, and engineering requirements necessary for efficient hypersonic flight, including combustion characteristics, engine performance, and thermal protection. Various hypersonic vehicles and their testing outcomes are reviewed to illustrate the potential and ongoing developments in this field.

1. Hypersonic Travel
A Game-Changing Technology
Dora E. Musielak, Ph.D.
University of Texas at Arlington - UTA Aerospace Professional Speaker Series
28 September 2016
Contact Author: dmusielak@uta.edu

2. Hypersonic Transport
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SST Mach 2.2 HST Mach 7

3. Hypersonic Missiles
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60 miles per minute
 Rapid response to time-critical targets.
 JP-fueled engine would enable a medium-size
rapid response ISR vehicle with operationally-
relevant range capability.
 Hypersonic missile offers advantage of flexible
recall and en route re-direction.
 Global fast-reaction reconnaissance missions.
Lockheed HSSW
X-51A

4. Hypersonic Air-Breathing Flight: Highlights
• Hypersonic Air-Breathing Propulsion (HAP)
Technology
– Dual-Mode Scramjet
• DARPA/USAF Military Programs
– From X-51A to HAWC and SR-72
• NASA Reusable SLV Concept: TSTO TBCC
• Europe: Skylon and HST
• Critical Technologies and Needed Breakthroughs
– Research Topics
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5. Aerothermodynamics-Propulsion
Aerothermodynamic performance cannot be decoupled from propulsion performance
due to shared surfaces and flow field interactions.
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Forebody Shaping, Shock Structure and Boundary-Layer State Impact Engine
Mass Capture and Combustor Performance
Variable Geometry
Dual-Mode Scramjet
Operation
Exhaust Plume and Nozzle Expansion Impact
Vehicle Pitching Moments and TPS
Exhaust Plume Interaction with
Airframe, Control Surfaces
Mach 7 CFD Model Image: NASA

6. Scramjet Propulsion 5 < M < 15
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Combustion
0
External
Compression
1
Internal
Compression
(Isolator)
3 4 8 9
Internal
Expansion
External
Expansion
ExpansionCompression
0  3 Adiabatic Compression
30 TT 
3  4 Isobaric heat addition
43 TT 
4  9 Adiabatic Expansion
94 TT 
Freestream
condition
Capture
area A0
0M 9AFuel injection
Air-breathing Engine  No Turbo-compressor
To produce static thrust, scramjet requires ram compression, i.e., forward flight speed
)( 090 VVmF  

7. Air-Breathing Propulsion Challenge
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Scramjet engine has a much more challenging task than rockets.
Consider flight at an optimum dynamic pressure:
2
2
0
0
V
q


Mass flow ingested by inlet
0
00
0000
2
V
Aq
VAm  
Inlet capture area A0 varies slightly with Mach number and angle of attack.
As cruise velocity increases , mass flow decreases  reducing thrust .
This is why scramjet engine has to occupy a large fraction of vehicle’s cross-section area.
Mach 10 vehicle needs ~ 80% of its frontal area to capture air
(subsonic airliner needs only 25%, and Mach 3 airplane uses about 40%).
)( 090 VVmF  

8. Air-breathing Propulsion Performance
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X-51
X-43A
f
AB
sp
mg
F
I
0
Specific Impulse
)s(spI
 090 VVmFAB  

9. Scramjet Critical Elements of Feasibility
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Compression and expansion processes are very
critical to overall propulsion system performance.
We estimate velocity ratios implied by Isp
fg
VV
Ma
fg
VV
V
fmg
F
Isp
0
09
00
0
09
0
00
1/1/ 




For M0 > 5 scramjets operate ~ stoichiometrically.
With H2 fuel f = 0.0293
sp
sp
I
MMa
gI
V
V
000
0
0
9 00096.00293.0
1 
0V 9V
0M
1
0
9

V
V
Fractional velocity change across engine is very small as M0 > 6.
A small inefficiency in nozzle or inlet could have large
consequences for overall scramjet engine performance.
Efficient Inlets and nozzles are essential for scramjet propulsion
Compression Expansion
 090 VVmF  

10. Fuel Energy
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Fuel is characterized by its energy content per unit mass.
Heat of reaction hPR represents (ideal) fuel energy density, i.e., fuel thermal energy per unit
mass of fuel.
Fuel hPR (BTU/lbm) hPR (kJ/kg)
Hydrogen, H2 51,571 119,954
Methane, CH4 21,502 50,010
Ethane, C2H6 20,416 47,484
Octane, C8H18 19,256 44,786
JP-4 18,400 42,798
JP-7 18,702 43,500 Rate at which chemical reactions make
energy available to engine cycle is
PRf hmrateenergyChemical
Overall efficiency of HAP cycle:
PRf hm
FV

0
0
rateenergyChemical
PowerThrust

Due to volume limitations, entirety of
available hPR cannot be realized. Burner
efficiency represents it
iPR
aPR
b
h
h
,
,

Heat addition in burner:
  PRbin hfhhq  34

)(HkJ/kg1014.1 2
5
 inq
7)-(JPkJ/kg1013.4 4
 inq

11. Fuel/Air Ratio, f
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Fuel/air ratio, indicator of combustion conditions
0m
m
f
f



General chemical equation for HCF + air indicates all carbon, hydrogen and oxygen atoms are
consumed in chemical reaction, yielding as products carbon dioxide and water:
22222
421
79
221
79
4
N
y
xOH
y
xCONO
y
xHC yx 


















 yx
yx
m
m
f
f
st



4103
336
0


For hydrogen fuel, x = 0 and y = 2, stoichiometric fuel/air ratio 0291.0stf
Ideal upper limit of fuel/air ratio is known as stoichiometric fuel/air ratio. It represents
condition where complete combustion of oxygen and fuel takes place.
For JP-7 fuel, x = 12.5 and y = 26, stoichiometric fuel/air ratio 06745.0stf

12. Specific Impulse of Hypersonic Propulsion
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 




















 111
2/1
00
0
0
t
in
KEsp
h
q
f
gf
V
I


Overall propulsion kinetic energy efficiency
KEnozzlerKEcombustoKEinletKE  0
Heat added per unit mass fuel PRbin hfq 
Curran, et al., 1991
Kinetic energy efficiency must be determined accounting for real gas effects and nonequilibrium
chemical reaction effects, very important in both combustors and nozzles.
KEOK 
Stoichiometric combustion
  PRbin hfhhq  34

)(HkJ/kg1014.1 2
5
 inq
7)-(JPkJ/kg1013.4 4
 inq 8M

13. Engine-Vehicle Integration
• Proper aerodynamic integration of propulsion system with vehicle is required for efficient
hypersonic flight. This includes integration of vehicle structures, material, thermal
management, controls, and subsystems.
• To capture required airflow, it is necessary to use entire forebody underneath vehicle as
compression surface, shaping it for high efficiency.
• Entire afterbody underneath vehicle must be used as a free expansion surface.
• Entire vehicle underside must be carefully designed to satisfy engine performance under all
flight conditions. Bottom of vehicle becomes an integral part of propulsion system.
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Shock
Shock
Compression
Engine
Forebody/Inlet Afterbody/Exhaust

14. X-30 National Aerospace
Plane (NASP) Studies
1987 - 1995
X-43A Hypersonic
Experimental Vehicle
16 Nov. 2004, it reached a cruise speed
of Mach 9.6 on H2 fuel.
Total flight time: 12 seconds
X-51A Waverider
On May 1, 2013, it reached a
cruise speed of Mach 5.1 on
HC fuel.
Total flight time: 210 seconds
Heritage HAP Technology
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HRE Program
1964 - 1974
Testing and Studies:
Australia HyShot
Russia CIAM
•GASL projectile
…….. .. more

15. NASA X-43A M10 Vehicle
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Goal: Demonstrate, validate and advance technology for hypersonic aircraft powered by an
airframe-integrated scramjet engine.
Flight History
-Flight 1: Mach 7 Target, June 2, 2001 Booster Failure
-Flight 2: Mach 6.91, March 27, 2004 Fully Successful
-Flight 3: Mach 9.68, Nov. 16, 2004 Fully Successful
First Flight of scramjet powered vehicle.
─ Demonstrated H2 burn, attained high
flight Mach number, non-symmetrical stage
separation (required for TSTO)
WEIGHT: 3,000 lbs / 1400 kg
Image: NASA

16. NASA X-43A Project
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Mach 9.65 Flight Nov 2004
Three-flight Project (scaled Mach 10 cruiser)
2 at Mach 7, power-on test 11 seconds*
1 at Mach 10
 Air launched on a highly modified Pegasus booster
 7 year project (1996 – 2004)
 ~ $230 M investment
 Second flight demonstrated scramjets can ignite at
Mach 6.83, fly at 94,000 ft (28.7 km) altitude.
Third flight maximum Mach 9.68, ~ 7 seconds
Image: NASA

17. X-51 WaveRider Mach 5
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Cruiser Length: 168 in ( 4.26 m), Max. Width: 23 in ( 0.58 m)
Stack Length: 301 in ( 7.6 m)
Pratt & Whitney Rocketdyne Scramjet
SJY61, Ethylene and JP-7 Fuel

18. X-51A WaveRider Flight Test
18
X-51A scramjet-powered waverider made
hypersonic history on 26 May 2010.
Vehicle soared at Mach 5 for over 3 minutes,
setting a new hypersonic world record.
Stack Length: 26 ft (7.6 m)
Empty weight: 4,000 lb (1,814 kg)
Scramjet: P&W Rocketdyne SJY61
Fuel: JP-7 (270 lbm)
Maximum speed: Mach 7+
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Scramjet can produce 400 to 1000 lb of thrust,
capable of accelerating more than 6437 km/h
(4000 mph)

19. X-51A Scramjet Engine Demonstrator
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Flight History
1st Flight – 26 May 2010: 143-second burn to Mach 5. Perfect flight would have lasted 100-sec more
and accelerated cruiser to Mach 6. Thermal seal breach at nozzle interface caused hot gases to leak.
2nd Flight – 13 June 2011: Booster to Mach 5. Scramjet lit, failed to transition to full power
3rd Flight – 14 August 2012: Fin Failure
4th Flight – 1 May 2013: 210-second combustion to Mach 5.1 and Parameter ID Maneuvers (Glide)

20. Scramjet Inlet Design
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21. Critical Aerothermodynamics
Viscous
Interaction
Aerodynamic
Heating
Shock Wave /
Boundary Layer
Viscous Interaction
(SWBLI)
Inlet Unstart
Inlet Spillage/BL
Ingestion
Bow Shock/Cowl
Shock Interaction
Jet
Interaction
Entropy Layer
Swallowing
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Hypersonic Aerodynamic environment is extremely harsh and hypersonic vehicles function near edge of
system capability.
Boundary Layer Transition

22. X-43A Coping with Aerodynamic Heating
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At M < 7, horizontal tail leading edges 2600°F (1400 °C) and nose 2000°F (1090 C°)
At Mach 10, Tw ~ 3600°F (1900°C)
Leading edge protection
Second X-43A vehicle, being mounted under right wing of B-52B
launch aircraft. Source: NASA DFRC Photo.

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Source: McClinton, 2006
Maximum working Temperatures
Titanium alloys: 1600R (890 K)
Superalloys: 1800R (1000 K)
Metal matrix composites: 4500R
Coated C-C composites: 5400R (3000 K)
Adiabatic wall temperatures in hypersonic flight can reach 5,000-15,000R (2778-8333 K)

24. TPS Cooling Requirements
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For thermal protection system we need to know whether existing fuel flow rate is sufficient to
satisfy total cooling requirements.
Total rate at which fuel can remove thermal energy
hfc is amount of thermal energy that can be absorbed by unit mass of fuel.
Assuming that most heat is due to turbulent flow, total rate at which heat must be removed
from engine surfaces (required cooling)
Ratio of available fuel cooling to required fuel cooling:
For a balanced system
Physical heat absorption capacities of hydrocarbon aerospace fuels crucial for cooling
fcff hmQ  
  0
5/1
0
5/11
0
05/1
0
2
000
A
A
q
V
mA
V
VV
AqQ w
wwwr
 




w
fc
r
f
A
A
V
q
K
hf
Q
Q 0
5/11
0
5/1
0



0
5/1
0
5/11
0
A
A
q
V
Khf w
fc 

25. Fuel Cooling Properties
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Physical heat absorbed by standard aerospace fuels in going from tank conditions to 1000 K (1800 °R). In all
cases, fuel is stored in liquid state and heated to vapor state before injecting into engine flowpath as a gas.
Hydrogen is best engine coolant for hypersonic vehicles. Costs and logistic complexities of
cryogenic hydrogen, as well as exceedingly low density force us to consider conventional
hydrocarbon fuels such as JP-7 for lower hypersonic flight regime.
JP-7 has better heat absorption than JP-8, and produces more of gaseous products that
reduce ignition times.
Fuel Liquid Density
lbm/ft3 (kg/m3)
fst Physical heat
absorption hfc
BTU/lbm (kJ/kg)
fsthfc
BTU/lbm (kJ/kg)
Hydrogen, H2 4.74 (75.9) 0.0291 6500 (15,100) 189 (440)
Methane, CH4 18.7 (300) 0.0583 1460 (3400) 85.1 (198)
JP-7, C12.5H26 49.5 (793) 0.0675 860 (2000) 58.0 (135)

26. Current HAP Military Programs
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 DARPA/USAF Hypersonic Air-breathing Weapon Concept (HAWC) flight demonstration of critical
technologies for effective and affordable air-launched hypersonic cruise missiles, hydrocarbon
scramjet-powered propulsion to enable sustained hypersonic cruise.
 Lockheed SR-72, unmanned, TBCC Mach 6 aircraft intended for intelligence, surveillance and
reconnaissance envisioned to succeed SR-71 Blackbird.
USAF/Lockheed Martin’s High Speed Strike Weapon (HSSW)
hypersonic missile. Image Credit: Lockheed Martin Co.
SR-72 twin-engine TBCC aircraft designed for a Mach 6 cruise.
Image Credit: Lockheed Martin Co.

27. Spaceplanes: Ground to Orbit
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28. Chemical Propulsion
28
Air Breathing
(Brayton Cycle)
Rockets
(Brayton Cycle)
Gas Turbine Engines No Rotor Engines
Turbo-
fan
Turbo-
jet
Pulse-
jet
Ramjet and
Scramjet
All Rockets
Turbofan +
AB
Turbojet +
AB
Hybrid or Combined Cycle Engine
Turbofan +
Ramjet
Turbojet +
Ramjet
Turbo-
Rocket
Ram-
Rocket
Rocket-
Scramjet
Pre-Cooled Air Cycle Engine Synergetic Air-Breathing and
Rocket Engine (SABRE)
Rocket
ee VmF 
)(0 oe VVmF  
Scramjet
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75% of weight at lift off
was propellant, and of
this, 83% was oxidant
(mLOX= 1.3 million lbm)

29. Combined Cycle Propulsion
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• TBCC
• Turbine-Based Combined Cycle (TBCC) use
turbine engines (TE) to boost the vehicle
to takeover (TO) speed.
• TEs are mounted in separate ducts to
facilitate cocooning to isolate them from
high enthalpy flow of hypersonic speeds.
TBCC uses high specific impulse (Isp) TE
booster for subsonic operation.
• Substantial data from testing, analysis,
and conceptual studies.
• RBCC
• Rocket-Based Combined Cycle (RBCC)
Engines have a rocket as primary propulsion
element to boost the vehicle to TO speed.
• Ejector rockets behave both as a rocket for
producing static thrust, and also as a ramjet
because inlet air is compressed and burned
with fuel to produce thrust.
• Fully integrated with one single flowpath.
• Limited data but substantial predictions
based on conceptual studies.

30. NASA Two-Stage to Orbit (TSTO) TBCC
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Booster: TBCC propulsion
Orbiter: all-rocket propulsion
0 < M < 3 Turbine w/rocket boost for takeoff
3 < M < 8 Dual-Mode Scramjet
8 < M < 10 for pull-up maneuver to staging point
(orbiter release)
cruise-back initially scramjet-powered with final
portion as glider to terminal landing.
NASA Image

31. NASA Baseline TSTO Mission Profile
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Unmanned, fully reusable and sized to deploy (or retrieve) a 20,000 lb payload to a 100100 nmi., 28.5° inclination orbit.
Ref: Snyder, Robinson, and Ferlemann, (2010)

32. TBCC Issues
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Issues:
 Dual integrated inlet operability and performance issues: unstart, inlet flow distortion
effects, bleed requirement/control, and operability margins
 Propulsion mode transition: high-temperature turbine  dual-mode scramjet
 Wide operating range  variable geometry  controls
 Integration/Interaction issues: Inlet/engine compatibility/unstart, etc.

33. SKYLON: SSTO Spaceplane
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Cylindrical payload bay: L = 13 m (42 ft 8 in) D = 4.8 m (15 ft 9 in)
Body: carbon-fiber-reinforced polymer space frame
Multiple layers of reflective foil thermal insulation
L = 82 m (269 ft), D = 6.3 m (21 ft)
LH2: 220 tonnes
GTOW = 275 tonnes
Source: REL Ltd.

34. Skylon Ascent Trajectory
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TO M ~ 0.5
M ~ 5.2
Karman line 100 km
Source: REL Ltd.

35. SABRE Rocket
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s800,2
0

gm
F
I
f
sp

Source: REL Ltd.

36. SABRE Rocket: Air Breathing Mode
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0 < M < 5.5
0 < h < 26 km (16 mi)
Air must be compressed to ~140 atm before injection into combustion chamber This raises air
temperature so high that it would melt material walls.
SABRE avoids this by first cooling air to −150 °C (−238 °F), using a pre-cooler heat exchanger until
air is almost a liquid. Then a conventional turbo compressor compresses air to required pressure.
Excess air is passed down a spill duct which has a
burner to recover some of drag losses.
Source: REL Ltd.

37. Hybrid Air-breathing Rocket
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A simple translating axisymmetric shock cone inlet slows air to subsonic speeds using two shock
reflections.
Some air passes through a precooler into central core, and remainder goes directly through a ring of
bypass ramjets.
Central core behind precooler uses a turbo-compressor run from same gaseous helium loop which
compresses air and feeds it into four high-pressure rocket combustion chambers.
Why a Curved nacelle? To help provide lift proportional to center of mass.

38. Civil Hypersonic Transport
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European Space Agency LAPCAT II: LAPCAT-A2 Mach 5
airliner powered by a pre-cooled air-turbo ramjet, and
Mach 8 airplane powered by a ramjet engine, both
fueled by LH2
JAPAN’s Exploration Space Agency: Hytex, Mach 5
hypersonic airliner propelled by a pre-cooled
turbojet engine
At Mach 5, air Tt ~ 1320K  impractical to compress
air due to work requirement and because resulting
compressor delivery temperature would be too high.
GTOM: 400 tonnes,
Fuel mass: 198 tonnes (LH2)

39. Issues for Hypersonic Airplanes
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Environment: Lifetime of water vapor decreases
exponentially, taking from 30 years at 25 km
altitude to less than one year above 32 to 34 km.
M8 LAPCAT 2 airliner will fly above 33 km to
minimize environmental impact.
Sonic Boom: generates 160-db noise that
travels to ground and permanently damage
human ears. Hypersonic planes to fly over
North Pole and cross Bering Strait, avoiding
populated land; it will fly higher so its ground
shock waves will be spread out and produce
a smaller shock wave
Aerodynamic Heating: Plane travelling at
M5+ will withstand surface temperatures
~1,000°C. Advanced materials required.
Fuel Consumption: M8 plane consumes two
times more LH2 per second than a M4 plane.
M8 plane gets to destination in half time so
fuel consumed overall is ~ same

40. Hypersonic Propulsion Challenges
• Operating efficiently and reliably over an
extraordinarily large range of flight conditions,
including, 0 < M < 6, 0 < M < 10, and from sea level
to orbital altitude.
• Accomplishing stable, efficient mixing and
combustion of fuel and air within burner of
reasonable size.
• Provide structural integrity necessary for reusable
system despite hostile environmental conditions.
• Integrate multiple cycle engines into a single
propulsion system capable of operating in multiple
modes, e.g., turbojet-ramjet-scramjet-rocket, or
rocket-scramjet-rocket.
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We are improving analytical tools, ground and flight testing facilities, and developing technologies
to finally prove that hypersonic air-breathing propulsion systems are ready for routine
operations.

41. Hypersonic Technologies
 Advanced High Temperature Materials - new
formulations and processes for metals and
composite materials
 Fuels and Injection Techniques
 Structures – lightweight primary structures,
reusable tanks for cryogenics
 Optimization of High Performance Scramjets
 Computational Fluid Dynamic (CFD) Methods
and Algorithms – computer simulation
 Heat Control – Cooling techniques and insulation,
regenerative fuels, etc.
 Ground and Flight Tests under realistic flight
conditions
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www.grc.nasa.gov/WWW/StructuresMaterials/AdvMet/research/titanium.html

42. Structures, Materials, and Thermal Management
• A scramjet-powered vehicle will take a more
leveled ascent trajectory.
• Due to low T/W scramjet needs more time to
accelerate. This depressed trajectory implies
that vehicle stays a long time in atmosphere at
hypersonic speeds, so atmospheric friction is a
huge problem.
• Heat addition produced by combustion at high
velocities and temperatures is another
significant factor.
• Materials for structure must have excellent
properties and be adequate to deal with these
and other hypersonic flow phenomena.
• Cooling of the engine’s structure by fuel or
radiation is essential.
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Glass, et al, 2002
At M < 7, X-43A’s horizontal tail leading edges 2600°F
(1400 °C) and nose 2000°F (1090 C°)
At Mach 10, Tw ~ 3600°F (1900°C)

43. U.S. Hypersonic Ground Test Facilities
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44. Hypersonic Scaling Laws
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What scaled propulsion components and initial conditions must be tested to provide
adequate scaling parameters with respect to various performance parameters of interest in
order to obtain a desired level of performance at a particular scale of the hypersonic engine?

45. Numerical Methods for Hypersonic Flows
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With limited ability to adequately represent hypersonic flow experimentally, hypersonic CFD
predictions becomes even more difficult because we lack substantial experimental data for a
variety of flow geometries and flight conditions.
Large-Eddy Simulation (LES) of HIFiRE scramjet
 Hypersonic flow dynamics is complicated by interaction of boundary layer and shear layer with shock waves,
leading to flow separation and instability not amenable to simple analysis. CFD aims to perform direct numerical
simulation (DNS) for all aerodynamic flows of interest.
 High-speed reacting turbulent flows are challenging to simulate fully. Large eddy simulation (LES) allows for
modeling of small scales of turbulence while resolving large-scale turbulent structures; currently limited to low
Reynolds number flows.

46. HAP Technologies and R&D Topics
• Dual-Mode Combustors
• Materials and Structures  light-weight,
high-T, high-strength O2 compatible, CC
• CFD Modeling, Simulation, and Prediction
• Ground Testing  diagnostics, scale, time
• Flight Testing  infrastructure, telemetry
• Thermal Management  TPS
• High-Energy Density Propellants
• TBCC Mach 4+ Turbine Acceleration
• Fuel Injection and Flameholding
• Design Optimization Methodology
• Hypersonic Aerothermodynamics
• Propulsion Flowpath Development
• Actuation Mechanism, Hot Valves, Seals
• TBCC Turbine/Scramjet Mode Transition
• TBCC Inlet Optimization
• RBCC Mode Transition Control
• Scaling Laws
• Combined and/or Hybrid Propulsion Cycles 
Dual-Mode Scramjet, Turbine/Scramjet/Rocket
• Design Optimization
• Boundary Layer Transition
• Combined Cycle Propulsion Mode Transition
• Airframe-Propulsion Integration 
Aerothermodynamic Performance
• Inlet Unstart Prediction/Control
• Combustor Flameout
• Shock Interaction Control
• Ignition/Flameout/Flameholding Testing,
Modeling and Simulation
• Combustion Stability
• Combustor/Isolator Interaction
• RBCC and TBCC Mode Transition
• RBCC Inlet/Nozzle Integration
• TBCC and RBCC Flow Path Optimization
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The Race for Hypersonic Travel Has
Begun …