Introduction to high speed propulsion musielak

Introduction to
High Speed Air-Breathing Propulsion (HAP)
Dora E. Musielak, Ph.D.
8 July 2017
All rights reserved. No part of this publication may be reproduced, distributed, or transmitted, unless for course participation, in any form or by any
means, or stored in a database or retrieval system, without the prior written permission of the Author. Contact D. E. Musielak, dmusielak@uta.edu

Introduction to High Speed Propulsion
• This lecture is intended to provide a top-level overview of high-speed air-breathing
propulsion and to provide a technical reference for the air-breathing propulsion systems
that are the focus of a 2-day course.
• In the course we highlight some unique challenges encountered in the design and build of
the advanced engines for future hypersonic vehicle applications.
• To begin the discussion, this presentatiom will focus on
– Types and Classification of Propulsion Systems
– Propulsion Performance Measures
– Thermodynamic Cycle Analysis
High-Speed Air-Breathing Propulsion Course
Dr. DORA E MUSIELAK
2
Mach 6 Air turbo-ramjet (ATR) from Aerojet
Mach 4.5 Tandem Turbo Ramjet from GE

High Speed Air Breathing Engines
Dr. DORA E MUSIELAK
3
Thrust: 23,770–29,160 lb (105.7 kN
with AB)
Bypass Ratio: 0.36
Overall Pressure Ratio: 32 to 1
Specific fuel consumption: Military
thrust: (0.73 lb/(lbf·h))
Thrust-to-weight ratio: 7.4:1
Pratt & Whitney F100-PW-100
or −220 AB turbofan
McDonnell Douglas F-15 Eagle → Mach 2.5
Mach 5.1 – X-51A Waverider
Mach 9.6 – X-43A
HC Scramjet
H2 scramjet

Chemical Propulsion
4
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 Engines
Turbofan +
Ramjet
Turbojet +
Ramjet
Turbofan+
Rocket+Ramjet
Rocket-
Scramjet
Pre-Cooled Air Cycle Engine Synergetic Air-Breathing and
Rocket Engine (SABRE)
Rocket
ee VmF 
)(0 oe VVmF  
Scramjet
Dr. DORA E MUSIELAK
Turbofan
Ram-
Rocket

Propulsion Design Considerations
Dr. DORA E MUSIELAK
5
Between Mach 3.0 to 4.0, gas entering engine has such a high temperature that no further energy
can be added without exceeding temperature limits of engine materials. No net thrust is possible.
If we can’t raise material temperature limits leaves us two possible design approaches:
(1) minimize or cancel turbomachinery work output requirements, or
(2) make turbomachinery work output independent of flight Mach number.
As flight Mach number increases, inlet diffuser compression ratio becomes high enough that
mechanical compression can be minimized or eliminated. Minimizing mechanical compression ratio
can lead to either a very low compression ratio, a variable mechanical compression process, or a
turbomachinery bypass process.
Complete elimination of mechanical compression leads to ramjet, but because a pure ramjet is not
effective at M < 3, we can combine ramjet with a turbo engine. This results in a combined cycle
engine  turbo-ramjet.
Having turbomachinery work output independent of Mach number leads to ATR engine.
At flight speeds > Mach 6.0, pressure losses in decelerating supersonic flow to subsonic speeds for
combustion are too high. This is overcome by burning fuel in a supersonic stream  scramjet

Mach 4.5 Tandem Turbo Ramjet (GE)
Dr. DORA E MUSIELAK
6
At takeoff and subsonic climb, core intake-guide-vanes are open, and engine operates as an
afterburning turbojet engine.
As engine climbs transonically and at low supersonic speeds, bypass begins to open allowing a
fraction of inlet flow to divert around turbomachinery and to mix with core flow before AB.
At Mach 4.5 cruise, intake guide vanes are closed, bypass is fully open, and engine operates as a
pure ramjet.

Mach 6.0 Air Turbo-Ramjet (ATR) (Aerojet)
Dr. DORA E MUSIELAK
7
Turbo-Ramjet includes air compression (ram + mechanical), constant pressure heat addition, and
expansion through a thrust nozzle.
Turbine is driven by high-temperature, fuel-rich gas from a separate gas generator (reaction
chamber).
After passing through turbine, this fuel-rich gas is mixed with airflow from turbocompressor and
burned in a combustor before expanding through nozzle.
Gas is formed by heating and vaporizing LH2 fuel in a dual-regenerator process. First heat
exchanger is located at turbine exit; second uses waste heat from combustor.

Mach 10.0 Turbo-Ram-Scramjet (P&W)
Dr. DORA E MUSIELAK
8
Initially, ram-scramjet is closed off
and turbojet provides thrust for
takeoff, climb, and acceleration to
transonic speeds.
Inlet geometry is varied to provide
for combined turbojet/ramjet
operation for initial supersonic
climb.
Between Mach 3.5 and 4 turbojet is
completely closed off, and engine
operates as a ramjet for climbing to
Mach 6.
From Mach 6 to Mach 10 cruise
altitude engine operates as a
scramjet.

Air-breathing Propulsion Performance
Dr. DORA E MUSIELAK
9
X-51A/HIFiRE-2
X-43A
)s(spI
Compare engine ability to produce thrust with a minimum of fuel expenditure
Ram/Scramjets
M0
   
0
0
0
00
0
,
fg
VV
gm
VVm
gm
F
I e
f
e
f
absp







000
,
g
V
gm
Vm
gm
F
I e
p
ep
p
rocketsp 



No one propulsion system is optimum over entire flight Mach number range

TriJet (Bulman and Siebenhaar, 2011)
Mach 6 Lockheed Martin
SR-72
From Take-Off to Hypersonic Flight
10
Dr. DORA E MUSIELAK
USAF/Lockheed Martin’s High Speed Strike Weapon
(HSSW) Mach 5+ hypersonic missile.
Mach 4 ZEHST
Zero Emission Hyper Sonic Transport

Turbo/Scramjet for Mach 6 SR-72
Dr. DORA E MUSIELAK
11
http://lockheedmartin.com/us/news/features/2013/sr-72.html

Propulsion Challenges
• Operation in several flight regimes poses many challenges for propulsion system
due to wide range of aerodynamic conditions.
• Disparities among hypersonic air-breathing propulsion flight requirements yield
difficult propulsion integration solutions.
• Integration and transition through multiple propulsion cycles is a huge issue.
• Closely coupling elements of various cycles: turbo-machinery, combustors (from
turbojets, ram/scramjets), gas generators (ejector rocket engines), heat
exchangers, air-breathing compression/inlet systems, and shared nozzles.
Dr. Dora E. Musielak
12
Low speed flight (Take-off  Mach 3)
Supersonic-low-hypersonic flight (Mach 3  Mach 6)
Atmospheric hypersonic flight (Mach 6  Mach 10)
SubOrbital/Orbital flight (Mach 10  Orbital speeds).
FlightRegimes

Technical Background
Dr. DORA E MUSIELAK
13
• Earth’s Atmosphere
• Ideal Gas Assumptions for Air
– Calorically perfect gas
– Thermally perfect gas
• Hypersonic Inviscid Flow Fields
– Euler Equations
• 1-D Aerothermodynamic Equations
– Total Enthalphy and Total Temperature
– Total Pressure
– Ideal Exit Flow Velocity and Mass Flow
– Impulse and Stream Thrust Function
• Constant Area Heating and Thermal
Chocking
• Shock Waves: Oblique Shocks, Normal
Shocks, and Expansion Flow Relations
In preliminary analysis we consider three topics:
channel flow  it provides an intuitive
understanding of propulsion flowpath;
shock waves  help us understand hypersonic
flows;
boundary layers  help us understand
complexity of hypersonic flow and scramjet
performance limitations .

Flight within Earth’s Atmosphere
Dr. DORA E MUSIELAK 14
Representative Atmospheric Properties
Hypersonic air-breathing flight occurs in stratosphere km)(52kft169Hkm)(11kft36 
K)(222R400,0

RT
)s/mN10(1.45s/ftlbf1003.3 2527
,0  
R
 Kms/(J101.98R)ftBTU/(s1018.3 26
,0   
Rk
m/s)(299ft/s980,0 Ra
Flight Mach number and flight velocity
M0 V0 (kft/s) V0 (km/s) V0 (mi/s)
1 0.980 0.2987 0.1856
1.02 1 0.3048 0.1894
3.348 3.281 1 0.6214
5.388 5.280 1.609 1
MIL-STD-210A
Standard day altitude H vs M0 contours for constant
dynamic pressure q0 ( from Heiser & Pratt, 1994).
2
2
0
0
V
q


If qo is too large, structural forces and drag on vehicle can be too high.
If qo is too small, surface area required for sustained flight may be too
large.

Air: Perfect Gas Assumption?
Dr. DORA E MUSIELAK
15
Engineering analysis of hypersonic air-breathing propulsion assumes air is perfect gas RTp 
Chemical composition of air: 79 % N and 21% O (by moles)
Equilibrium static enthalpy

p
eh 
Equilibrium specific heat s
p
p
T
h
C 








Air behaves as calorically perfect gas:
K)(400R720K)(217R390 
T
Air behaves as thermally perfect gas:
K)(1700R3000K)(400R720 
T
Equilibrium ratio of specific heats
v
p
c
c

v
v
T
e
C 








For calorically perfect gas
For thermally perfect gas
40.1
286.1

Freestream Mass Flow
Dr. DORA E MUSIELAK
16
Specific Thrust
Hypersonic air-breathing engines generate thrust in direct proportion to rate at which they are
able to capture and process surrounding atmosphere.
Freestream Mass Flow per Unit Area
0q
Standard day altitude H vs M0 contours for constant
freestream mass flow per unit area (Heiser & Pratt).
Engine’s total uninstalled (internal) thrust is
proportional to total mass flow rate of air
ingested.
0000 AVm 
00000 MaV  
00
0
const
a
a
M SLSL
SL




00
0
0
0
00
22
Ma
q
V
q
V 
Air mass flow entering engine
Freestream mass flow per unit area
Flight Mach number at any
altitude
For high M0, vehicle needs very large capture area!
 090 VVmF  

Thermodynamic Cycle Analysis
Dr. DORA E MUSIELAK
17
The goal of this section is to give an overview of the ideal performance of high speed air
breathing engines based on thermodynamic performance.
The thermodynamic cycle analysis models a semi- ideal heat engine based on the closed
Brayton cycle.
We typically use the Stream Thrust Analysis method. The method relies heavily on momentum
relationships.
For preliminary performance analysis of hypersonic airbreathing engine the Stream Thrust
Analysis method is preferred because it takes into consideration mass addition, momentum
and kinetic energy fluxes contributed by the fuel, and it helps us perform more extensive
parametric analysis .

Engine Reference Stations
Dr. DORA E MUSIELAK
18
Combustion
0
External
Compression
1
Internal
Compression
(Isolator)
3 4 9 10
Internal
Expansion
External
Expansion
ExpansionCompression
0  3 Adiabatic Compression 30 TT 
3  4 Isobaric heat addition 43 TT 
4  10 Adiabatic Expansion 104 TT 
Freestream Static Temperature0T
3T Burner Inlet Static Temperature
Burner Exit Static Temperature
4T
Freestream
condition
Capture
area A0
0M 10AFuel injection

Thermodynamic Cycle Analysis
Dr. DORA E MUSIELAK
19
   01034 hhhhWnet 
 34 hhQadded 
 010 hhQreject 
 
10
0
4
3
TdsTdsWnet

Isobaric Process 3  4
Isobaric Process 10  0
Brayton cycle efficiency (ideal)
03 /
1
1
TTQ
W
added
net
tc  


Net power output
Cycle static temperature ratio
0
3
T
T

t4 Thermal ceiling

Ideal Thermodynamic Cycle Efficiency
Dr. DORA E MUSIELAK
20
34
010
1
hh
hh
Q
W
added
net
tc


 


)kJ/kg77(BTU/lbm330 h
)kJ/kg1423(BTU/lbm6123 h
549.0tc High overall engine efficiency 0
ptc  0
powermechanicalEngine
powerThrust
efficiencypropulsive p
Representative
values
K)(1556R28003

T

Engine Cycle and Parameter Trends (M0 < 5)
Dr. DORA E MUSIELAK
21
COMPRESSION COMBUSTION EXPANSION
3 40
Burner exit total temperature Tt4
represents cycle temperature limit
3
3
3
hth =1-
T0
T3
Tt3 /T0 ~ 3
T3 < 1670 K

Thermodynamic Properties
Dr. DORA E MUSIELAK
22
For T > 3000R (1700 K), Cp depends
strongly on both T and p. Values of cp at
higher T are due to dissociation of O2.
Burner entry temperature must be
controlled due to air dissociation during
heat addition process.
Any internal molecular energy invested
in dissociation will likely be lost or
unavailable for exhaust flow kinetic
energy, and thus reduce
thermodynamic cycle efficiency.
A more complex phenomenon arises
when energy addition is due to actual
combustion kinetics.Maximum Allowable Compression Temperature T3: 2600 –
3000R (1440 – 1670 K)  T3 avg = 2800R (1560 K)
Avg. Cp for heat
addition process  
RBTU/lbm396.0~
/ln 34
34



TT
ss
Cp

Engine Inlet Compression Process
Dr. DORA E MUSIELAK
23
 Compressor flow may be considered adiabatic
 Heat transfer at engine inlet may be neglected
 Viscous dissipation in wall boundary layer and
shocks account for sources of irreversibility
 Compressor efficiency: adiabatic compressor
efficiency 𝜂c and polytropic efficiency (both are
interrelated).
 Compressor adiabatic efficiency is a function of total
pressure ratio 𝜋c and decreases with as pressure
ratio increases
23 / ttc TT

 /1
 cc
r
d
t
t
d
M
p
p




 )1/(
2
0
0
2 2
1
1




 


1
1/1




c
c
c




 
2/
2/
2
0
2
2
02
02
V
V
hh
hh ideal
t
st
d 



pc º pt3 / pt2

Adiabatic Compression Efficiency
Dr. DORA E MUSIELAK
24
Adiabatic compression efficiency is strongly dependent upon number of oblique shocks.
Inlet must have three or four oblique shocks in order to achieve needed air compression.
Assumes no friction forces or aerodynamic heating

Ram Pressure Recovery
Dr. DORA E MUSIELAK
25
Ratio between stagnation pressure in Station 3 and Station 0, πd, gives loss in total pressure
associated with compression process. Total pressure ratio is heavily influenced by shock wave-
boundary layer interactions and by viscous loss as flow stagnates due to no-slip condition at wall
surface.
total pressure recovery exponentially decreases with M0

Inlet Kinetic Efficiency
Dr. DORA E MUSIELAK
26
M. Smart’s “Scramjet Inlets” RTO-EN-AVT-185
Methods for determining properties
at inlet throat:
(1) use an empirical relation for ηKE
or ηKE_AD in combination with a
capability parameter (see below)
(2) use CFD to perform a numerical
simulation of forebody/inlet
flowfield.
Empirical relation by Waltrup:
For higher compression inlets this
correlation is conservative.

Maximum Allowable Compression Temperature
Dr. DORA E MUSIELAK
27
T3 must be limited to value that prevents excessive dissociation in exhaust flow.
Maximum allowable burner entry temperature T3 requires elaborate analysis and computations as
this temperature depends on many interrelated variables, including flight altitude, M0, inlet losses,
fuel type, f, burner and nozzle geometry.
Maximum allowable burner entry
temperature:
1440 K < T3 < 1670 K
2600°R < T3 < 3000°R
Entire adiabatic compression process will take place where air behaves as a thermally perfect
gas and dissociation effects are negligible.
3
t3
Tt4 = 2200K

Required Burner Entry Mach Number
Dr. DORA E MUSIELAK
28
Stagnation Temperature of Inlet Flow





 





 
 2
33
2
00
2
1
1
2
1
1 MTMTT cc
t

Burner Entry Mach Number











 


 1
2
1
1
1
2 2
0
3
0
3 M
T
T
M c
c


Straight lines are hypersonic asymptotic limit.











 

 1
2
1
1
2
0
3
0
T
T
M c
c


3
0
0
3
T
T
M
M
 38.0
2800
400
0
3

M
M
when
13 M
We need
so T3< T3max allowable

Combustion Process
Dr. DORA E MUSIELAK
29
▪ Fuel is characterized by its heating value hPR
(maximum releasable thermal energy per unit
mass)
▪ Burner is characterized by its efficiency 𝜂b, and its
total pressure ratio 𝜋b
▪ Sources of irreversibilities (burner loss) are
burning at finite Mach number, frictional losses on
walls and turbulent mixing
▪ Thrust control/engine design parameters are Fuel-
to-air ratio f and burner exit temperature Tt4
▪ Application of energy balance across burner yields
either f or Tt4.
    404030 1 ttfbPRft hfmhmmhmhm   
4
34
tbPR
tt
hh
hh
f




 
f
cfhTcc
T
pbPRtpp
t



1
// 4343
4

f =
tl -trtc
hPRhb / h0 -tl
tl =
ht4
h0
hb º
hPR,actual
hPR,ideal
Burner Temperature Limit
1
3
4

t
t
b
p
p

Tt4 = 2021K (TO)P&W 4098 Turbofan

Fuel Heat of Reaction
Dr. DORA E MUSIELAK
30
Fuel is characterized by its energy content per unit mass.
Heat of reaction or heating value 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 mainly 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

Fuel/Air Ratio
Dr. DORA E MUSIELAK
31
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 stoichiometric fuel/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

Equivalence Ratio
Dr. DORA E MUSIELAK
32
Stoichiometric proportion of fuel to oxidizer results in neither excess oxygen nor any excess fuel.
Any more fuel would result in unburned fuel in products of combustion, and any more air would
result in excess oxygen in products. Nitrogen is treated as remaining unreacted (or inert in
chemical terms) in combustion process.
Equivalence ratio 𝜙 describes fuel lean or rich condition of a combustor
stf
f

Highest combustion temperature is achieved very near stoichiometric ratio.

Effective Fuel/Air Ratio
Dr. DORA E MUSIELAK
33
T-S diagram for ram/scramjet, from H&P
book. Constant-pressure heat addition
and rejection for H2-air combustion.
549.01 4
3
10
0



Tds
Tds
tc
Heat added per unit mass of air:
3
4
3
4
0
hhdsThf
m
hm
PRb
PRfb
 



Combustion efficiency accounts for
incomplete combustion
For H2+air, effective fuel/air ratio
0263.0
4
3


PR
b
h
Tds
f
More complex phenomena arise when energy
addition is due to actual combustion (chemical
reaction modeling required).
hb = combustion efficiency

Fuel/Air Ratio and Flammability
Dr. DORA E MUSIELAK 34
Equivalence ratio of 0.52 presents a
problem as combustion is near lean
flammability limit
Overcome by burning a local rich
fuel/air mixture and/or stabilizing
flame …
Attaining stable combustion
ensures complete
combustion for lean mixtures
(Kerosene-Type Fuel in Air)
… we still cannot predict how lean - before beginning of instability!
06745.0stf

Total Pressure Loss in Combustor
Dr. DORA E MUSIELAK
35
Combustion causes a loss in total pressure, and a rise in total temperature
Isobars at combustor entrance pt3
and exit pt4 show a total pressure
drop:
  34 ttburnert ppp 
 burnertpin Tcmq  0

Thermal power input (by fuel) is
proportional to temperature rise
across combustor:
Combustor sources of irreversibilities: Burning at finite Mach number,
frictional and thermal losses on walls, turbulent mixing, …

Expansion Process
Dr. DORA E MUSIELAK
36
▪ Nozzle primary function: accelerate gas efficiently
▪ Gross thrust parameter Fg gives nozzle’s
contribution to engine thrust
▪ Gross thrust reaches a maximum when nozzle is
perfectly expanded: p9 = p0
▪ Real nozzle flows may be considered adiabatic
▪ Nozzle losses manifest themselves as total
pressure loss
▪ Imperfect nozzle expansion is caused by a
mismatch between nozzle area ratio and altitude
of operation
▪ Underexpansion is caused by smaller-than-
necessary nozzle area ratio, leading to p9 > p0
▪ Overexpansion is caused by larger-than-necessary
nozzle area ratio, leading to p9 < p0.
  90999 AppVmFg  
1
/)1(
9
7
/)1(
/)1(
9
7












 







p
p
p
p
t
n
t
n

Air-Breathing Propulsion Performance
Dr. DORA E MUSIELAK
37
Performance parameters  figures of merit useful from supersonic to hypersonic flight.
Assumption: 1-D model; exhaust flow is perfectly expanded to surrounding atmospheric pressure, a
condition we attempt to attain because it maximizes thrust.
Specific Thrust
Specific fuel consumption
Specific Impulse
Overall Efficiency
0
00

V
hf
m
F PR

F
m
S
f


0
000

Vg
h
mg
F
I PR
f
sp

sp
PRPRf
I
h
Vg
hm
FV
 000
0


 
  















22
1
22
1
2
0
2
0
2
0
2
0
VV
fm
FV
fh
VV
f
e
f
PR
e
pth



Air-Breathing Propulsion Performance
Dr. DORA E MUSIELAK
38
Thermal Efficiency
Propulsive Efficiency
Overall Efficiency
Engine overall efficiency increases as flight speed increases, and it approaches thermal
efficiency. Scramjet offers competitive performance for hypersonic flight!
hth =
1
2 V10
2
-V0
2
( )
fhPR
=
V10
2
V0
2
-1
fhPR
V0
2
/ 2
=hb ×htc
hp =
2
V10
2
V0
2
+1
=
2
hth ×
fhPR
V0
2
/ 2
+1+1
ho =
2
V10
2
V0
2
+1
æ
è
ç
ö
ø
÷
fhPR
V0
2
/ 2
=
2 hth ×
fhPR
V10
2
/ 2
+1 -1
æ
è
ç
ö
ø
÷
fhPR
V0
2
/ 2
ratiofuel/airf
reactionofheatPRh
PR
o
o
hf
V
V
V 2
0
10
1






Energy made available by
chemical reaction/kinetic
energy of freestream air
(H2)kJ/kg3492PRhf
K1556K,222 30  TT
atm50.2atm,01.0 30  pp
Heiser&Pratt

Cycle Static Temperature Ratio
Dr. DORA E MUSIELAK
39
0
3
T
T

Cycle static temperature ratio is a
principal factor in thermodynamic
cycle efficiency and can be used to
impose limit of maximum allowable
compression temperature.
It also influences overall efficiency.
As shown in plot, there is no gain by
increasing  indefinitely.
Maximum cycle efficiency when
001 TC
hf
p
PRb
ec
ec 


 


75.7
90.0 ebc 
1.14
00

TC
hf
p
PRb
ho =
2 hth ×
fhPR
V10
2
/ 2
+1 -1
æ
è
ç
ö
ø
÷
fhPR
V10
2
/ 2
y =
T3
T0
60 M
100 M
Heiser&Pratt

Specific Impulse of Hypersonic Propulsion
Dr. DORA E MUSIELAK
40
 




















 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

Ramjet Engine
2 < M0 < 5
Diffuser/Inlet Throat
Normal
Shock
Fuel Injection Subsonic
Combustion
Nozzle
Cowl Flame Holders
M =1Centerbody
Air
Ram effect: when volume of air is forced into small space at high enough speeds, it is
compressed to a higher pressure.
Air passage through one or more shocks slows down, compresses, and heats air flow.
41
Dr. DORA E MUSIELAK
M < 1
Ramjet: cannot deliver thrust without forward motion. An auxiliary device needed to accelerate
ramjet to speeds at which it can provide forward net thrust.
1M
10 M

Ramjet Performance: Optimum M0
Dr. DORA E MUSIELAK
42
4tT  1
1
2 3/1
/max0 0


 

mFM 
0
4
T
Tt

By using SUPERSONIC COMBUSTION,
temperature rise and pressure loss due to
deceleration through inlet can be reduced.
At M0 > 6, high static temperatures reduce
ramjet performance.
Inlet total pressure recovery exponentially
decreases with M0
Pressure and temperature ratios would be
unfavorably high if engine continued to
operate as a subsonic combustion ramjet.
1) High degree of dissociation of
combustor exhaust flow, reduce energy
available for exhaust velocity.
2) Pressures far too high for Brayton cycle
operations or to withstand by structure. At M0 > 6 gas dissociation limits Tt4

Scramjet Critical Elements of Feasibility
Dr. DORA E MUSIELAK
43
Compression and expansion processes are very
critical to overall propulsion system performance.
We illustrate this by estimating velocity ratios that
are implied by specific impulse
fg
VV
Ma
fg
VV
V
fmg
F
Isp
0
09
00
0
09
0
00
1/1/ 




Above M0 5 scramjets operate approximately
stoichiometrically. For H2 fuel f = 0.0293 ~ const
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 for M0 > 6.
A small inefficiency in nozzle or inlet could have large
consequences for overall scramjet engine performance.
High efficiency inlets and nozzles are essential for scramjet
propulsion
Compression Expansion
 090 VVmF  

Hypersonic Air Breathing Propulsion (HAP)
• Ram/Scramjet Propulsion
• Main Scramjet Engine Components
– Inlet-Isolator
– Combustor
– Nozzle
• Engine-Vehicle Integration
• Hypersonic Propulsion Challenges
• Ground Testing
• CFD and Numerical Methods applied to HAP
• Technology Issues and Critical Design Issues
• X-43A and X-51A Aircraft Development
• The Future for HAP
• U.S. Technology Development Roadmaps
44
Dr. DORA E MUSIELAK
The 2-day course consists of 16 one-hour sessions, which
allows participants to study in detail the following topics:

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 altitudes.
• 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.
Dr. DORA E MUSIELAK
45
We must develop analytical tools, ground and flight
testing facilities, and prove that hypersonic
propulsion systems are ready for routine operations.

Technology Issues
▪ 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
46
Dr. DORA E MUSIELAK
www.grc.nasa.gov/WWW/StructuresMaterials/AdvMet/research/titanium.html

Hypersonic Flow Regime
Hypersonic flow regime is where most
of the flow total temperature exists as
kinetic energy.
Mach number changes because static
temperature and speed of sound are
changing.
20
0
2
2
1
2
M
Tc
Tc
Tc
V
H
K
p
tp
tp



91.083.0  K
47
Dr. DORA E MUSIELAK
0tT = stagnation temperature of freestream flow
Technical differences between air-breathing
engines operating at hypersonic speeds and
those for lower speeds stem mainly from high
stagnation temperature levels at high M0
  2
02
1
00 11 MTTt  
V > 1.5 km/s

Scramjet Engine: 5 < M < 15
Simple concept, yet scramjet components are quite complex in their design and operation.
48
Dr. DORA E MUSIELAK
W. Engelund, NASA Langley, May 2001
M > > 1

Air-Breathing Propulsion Challenge
Dr. DORA E MUSIELAK
49
Scramjet engine has a much more challenging task than rockets.
Let us consider a flight at an optimum dynamic pressure:
2
2
0
0
V
q


Mass flow ingested by inlet
0
0
000
2
V
Aq
AVm  
where A is inlet capture area, which varies slightly with Mach number and angle of attack.
As cruise velocity increases , mass flow decreases  reducing thrust .
That 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%).

Tip-to-Tail Propulsion Cycle Analysis
Dr. DORA E MUSIELAK
50
SCRAM performs nose-to-tail simulation of real gas flow with equilibrium thermodynamic
characteristics encountered in hydrogen-fueled ramjet/scramjet engine. Written for supersonic
flows, code modified to handle subsonic flows and dual-mode combustor operation.
SEAGULL (2D/3D Euler)
Forebody/inlet shock losses
SCRAM (1D) with EQ chemistry
Combustor cycle analysis (CV process)
HUD (Boundary Layer)
Forebody/inlet/combustor/nozzle
Heat and friction losses
SEAGULL (2D/3D Euler)
Nozzle expansion losses
SRGULL code – NASA hybrid program comprised of codes SCRAM and SEAGUL.
Hypersonic propulsion modeling and analysis code developed by combining a
combustor program and a flow-field analysis program.

Fastest ABP Vehicles
51
Dr. DORA E MUSIELAK
M 6-8 – In 2012, HIFiRE-2 was
first flight test of hydrocarbon-
fueled scramjet, accelerating
from Mach 6 to Mach 8,
launched by 3-stage sounding
rocket system.
M 9.6 – X-43A achieved Highest
Hypersonic Speed with H2
Fueled Scramjet.
On 16 Nov 2004, vehicle cruised
at Mach 9.6 for 20 seconds.
X-43A/Hyper-X program
provided first free flight data on
scramjet engines,
demonstrating predictive
design tools were accurate.
M 5.1 – X-51A achieved
Highest Hypersonic Speed
with HC Fueled Scramjet.
On 1 May 2013 final flight,
vehicle reached Mach 5.1
traveling more than 230 nm
in just over 6 min. Longest of
four X-51A test flights and
longest air-breathing
hypersonic flight on HC.

X-43A Project Overview
Dr. DORA E MUSIELAK
52
▪Three-flight Project
▪2 at Mach 7, power-on test 11 seconds*
▪1 at Mach 10
▪ Scaled version of a Mach 10 cruise configuration
▪ Air launched on a highly modified Pegasus
booster
▪ 7 year project (1996 – 2004)
▪ ~ $230 M investment
Second X-43A flight demonstrated that
scramjets can work at conditions up to
Mach 6.83 at 94,000 ft (28.7 km)
altitude.
Third flight, demonstrated maximum
powered Mach 9.68, ~ 7 seconds
During scramjet operation vehicle
achieved cruise condition, F = D

NASA X-43A Program
Dr. DORA E MUSIELAK
53
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)

X-51A WaveRider Flight Test
54
The X-51A scramjet-powered waverider
made hypersonic history on May 26, 2010.
The 4.26 m long demonstrator vehicle
soared at Mach 5 for over 3 minutes, setting
a new hypersonic world record.
270 lb. of JP-7
Stack Length: 26 ft (7.6 m)
Empty weight: 4,000 lb (1,814 kg)
Scramjet: P&W Rocketdyne SJY61
Fuel: JP-7
Maximum speed: Mach 7+
Dr. DORA E MUSIELAK
Scramjet can produce 400 to 1000 lb of thrust,
capable of accelerating more than 6437 km/h
(4000 mph)

Hypersonic Flight Critical Design Issues
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
55
Dr. DORA E MUSIELAK
Hypersonic Aerodynamic environment is extremely harsh and hypersonic vehicles function near edge of
system capability.
Boundary Layer Transition

Skin Temperature of Hypersonic Vehicle
Dr. DORA E MUSIELAK
56
Due to extreme airstream total temperature, external vehicle skin and exhaust nozzle of X-51A
vehicle (4.5 < M < 6.5) experienced temperatures ranging from 1500F to 3500F (1090-2200 K).
Vehicle’s external TPS and nozzle liner are a combination of ablative thermal protection materials
and third generation reusable surface insulation (Shuttle tile) material. Tiles are used in regions
where shape stability is required such as inlet ramp and cruiser windward surfaces.
Highest heat fluxes on X-43A required
installation of C-C composite located at
nose leading edges, horizontal control
surfaces and tail leading edges.
(Ohlhorst et al. 2006).
2
0
0
0
2
1
1 M
T
Tt 


Vehicle size: 12 ft x 5 ft
Nose leading edge Tmax
- Mach 7, < 3000F
- Mach 10, ~ 4000F
C-C side chims
C-C nose leading edge
Upper surface of vehicle
Tail leading edges
-Mach 7, Haynes alloy
- Mach 10, C-C
C-C horizontal
control surface
Hypersonic aerodynamics  In stagnation flow
regions temperatures reach values that exceed
durability limits of materials:

Numerical Methods
Dr. DORA E MUSIELAK
57
With limited ability to adequately represent hypersonic flow experimentally, hypersonic CFD
predictions becomes even more difficult because substantial experimental data for a variety of
flows and flight conditions are not available.
Hyper-X vehicle at M∞ = 7
Large-Eddy Simulation (LES) of HIFiRE scramjet
Fluid dynamics of hypersonic flows is complicated by interaction of boundary layer and shear layer with shock
waves, leading to flow separation and instability not amenable to simple analysis. Goal of CFD 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, but is currently limited to
low Reynolds number flows.

Nomenclature
Dr. DORA E MUSIELAK
58

References
• Heiser W.H. and Pratt, D.T., Hypersonic Airbreathing Propulsion. AIAA Education Series, 1995
• Bulman, M.J. and Siebenhaar, A., “Combined Cycle Propulsion: Aerojet Innovations for Practical Hypersonic Vehicles.” 17th AIAA
International Space Planes and Hypersonic Systems and Technologies Conference, 11 - 14 April 2011, San Francisco, CA, AIAA 2011-
2397.
• Anderson, J. Jr, Hypersonic and High Temperature Gas Dynamics. AIAA Education Series, 2006.
• Curran, E.T., et al., THE USE OF STREAM THRUST CONCEPTS FOR THE APPROXIMATE EVALUATION OF HYPERSONIC RAMJET ENGINE
PERFORMANCE, AD-769 481, Air Force Propulsion Lab, 1973.
• Smart, M., “Scramjet Isolators,” RTO-EN-AVT-185
• Curran, E.T., Leingang, J.L, Carriero, L.R., and Petters, D.P., “A Review of Kinetic Energy Methods in High Speed Engine Cycle Analysis,”
International Symposium on Air Breathing Engines, Paper ISABE 91-10.5(L), AIAA 1991
• Kerrebrock, J.L., Aircraft Enginesand Gas Turbines, MIT Press, 2001
• Mattingly, J.D., Elements of Propulsion: Gas Turbines and Rockets, AIAA Education Series
• Mutzman, R. and Murphy, S., “X-51 Development: A Chief Engineer’s Perspective.” 17th AIAA International Space Planes and
Hypersonic Systems and Technologies Conference. 13 April 2011.
• Marshall, L.A., Corpening, G.P., and Sherill, R.,“A Chief Engineer's View of the NASA X-43A Scramjet Flight Test,” May 2005.
• X-51A Fact Sheet, www.af.mil/AboutUs/FactSheets/Display/tabid/224/Article/104467/x-51a-waverider.aspx
• Musielak, D.E., Advanced High Speed Propulsion Technologies, AIAA Short Course Lecture Notes, March 2017
Dr. DORA E MUSIELAK
59
If you wish more information about the complete 2-day course, please
email me dmusielak@uta.edu

Disclaimer
Dr. DORA E MUSIELAK
60
The material presented in this Lecture is based on knowledge freely available and published in
books, technical reports, public briefings, conference papers and journal articles. A list of references
is provided to facilitate further examination and extension of material.
Because we face very strict compliance requirements in the United States on ITAR (International
Traffic in Arms Regulation), this limits the availability of some engine performance parameters and
data. All images used to illustrate hypersonic technologies are available in the Internet.
The views expressed in this Lecture are those of the Author and do not reflect the official
policy or position of the United States Government, AIAA, or any other institution mentioned.
All rights reserved. No part of this course may be reproduced, distributed, or transmitted
in any form or by any means, or stored in a database or retrieval system, without the
prior written permission of the Author. For permission to use this material, please
contact D. E. Musielak, dmusielak@uta.edu

Dora E. Musielak, Ph.D.
Dr. DORA E MUSIELAK
61
Dr. Dora E. Musielak has directed government sponsored R&D projects in industry and
academia. Her key expertise is in high-speed air breathing propulsion and liquid chemical
rockets.
As chief scientist, Dr. Musielak led a scramjet propulsion development program sponsored by
the U.S. DoD. She has authored numerous reports and papers related to high speed propulsion
(scramjets, rockets, and detonation engines), with focus on numerical simulation of fuel
injection, high speed reacting and nonreacting turbulent flows.
Dr. Dora Musielak is the recipient of two NASA research fellowships, one of which she was
awarded to carry out research at the Hypersonics Propulsion Branch, NASA Langley Research
Center. At NASA, Musielak began research related to scramjet combustion, including modeling
and simulation of fuel injection, mixing, and flameholding using the VULCAN code.
An AIAA Associate Fellow, Dr. Musielak is a research professor at the University of Texas at
Arlington. Musielak has served in several national technical committees, including the NRC
Committee on Breakthrough Technology for Commercial Supersonic Aircraft, the AIAA Pressure
Gain Combustion Program Committee (PGC PC), and the AIAA High Speed Air Breathing
Propulsion TC, a committee she chaired from 2014 to 2016.

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