Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

309
The system operates by keeping a constant pressure in chamber 10, equal to the preset value
(proportional to the spring 16 pre-compression, set by the adjuster bolt 15). The engine’s
necessary fuel flow rate
i
Q and, consequently, the engine’s speed n, are controlled by the
co-relation between the
c
p
pressure’s value and the dosage valve’s variable slot
(proportional to the lever’s angular displacement

).
An operational block diagram of the control system is presented in figure 4. TURBO-JET
ENGINE
FUEL
INJECTION
DOSAGE
VALVE
FUEL
INJECTION
PUM P
PRESSURE
CONTROL LER
(

Fig. 4. Constant pressure chamber controller’s operational block diagram
3.2 System mathematical model
The mathematical model consists of the motion equations for each sub-system, as follows:
a.
fuel pump flow rate equation

(,)
pp
QQn
y

, (5)
b.
constant pressure chamber equation

ipA
QQQ

 , (6)
c.
fuel pump actuator equations

2
2
4
A
AdA cA
d
Qpp


t

  , (9)
d.
pressure sensor equations

0
2
()
snn A
Qdzxpp


   , (10)

2
12 2
()
4
n
mc A e
d
lS
p
l
p
lk z x


, (11)

, (13)
where , , ,
p
iAs
QQQQ are fuel flow rates,
c
p
-pump’s chamber’s pressure,
A
p
- actuator’s A
chamber’s pressure,
CA
p
-combustor’s internal pressure,
0
p
-low pressure’s circuit’s
pressure,
dA

,
n

,
i

-flow rate co-efficient,
,
An

T
-engine’s
inlet’s parameters (total pressure and total temperature).
It’s obviously, the above-presented equations are non-linear and, in order to use them for
system’s studying, one has to transform them into linear equations.
Assuming the small-disturbances hypothesis, one can obtain a linear form of the model; so,
assuming that each X parameter can be expressed as

 
2
0

1! 2! !
n
XX
X
XX
n


  
, (14)
(where
0
X
is the steady state regime’s X-value and
X

-deviation or static error) and
neglecting the terms which contains

kxkz

  
, (17)

ipA
QQQ

 
, (18)

0
dd
dd
AsA AA
QQV
p
S
y
tt


    
, (19)


1
2
emc
l


Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

311
where the above used annotations are

2
10
00
2
2
,
8
ic
A
AdA i
cA
b
p
d
kk
pp









d
p
kk
p
   
 


. (22)
Using, also, the generic annotation
0
X
X
X

 , the above-determined mathematical model can
be transformed in a non-dimensional one. After applying the Laplace transformer, one
obtains the non-dimensional linearised mathematical model, as follows



s1 s
PA A A
y
cx cz c
k
py
kx kz
p


, (26)

cQ c Q i
kp k Q


. (27)
For the complete control system determination, the fuel pump equation (for
p
Q
) and the jet
engine equation for
n (Stoicescu & Rotaru, 1999) must be added. One has considered that the
engine is a single-jet single-spool one and its fuel pump is spinned by its shaft; therefore, the
linearised non-dimensional mathematical model (equations 23÷27) should be completed by

pp
n
py
Qknk
y
, (28)


*
1
s1
MciHV
nkQ k p


20
,, , ,
2
smc
i
yzxc
eye e AA
kS p
k
ml
Tkk
kTk klkp





  

0
0
0
00 0
,,,
p
Aic
ic c


are becoming
null. The fuel flow rate through the actuator
A
Q is very small, comparative to the
combustor’s fuel flow rate
i
Q
, so
p
i
QQ

. Consequently, the new, simplified, mathematical
model equations are:
-
for the pressure sensor:

lc z
xkp kz, (31)
where

00
0
1
0020
0
0
,
cc

lc
p
kz
pz
k
p
z






, one obtains



lci c
xk
pp
  ; (33)
-
for the actuator:



s1
y
x
y


 

 
 




 

 
,

0
00
0
0
0
2
0
0
00
4
2
s
nnA f
x
s
A



*
1
s1
MciHV
nkQ k
p

 
, (36)

i
pp
n
py
QQ knk
y
  , (37)



lci c
xk
pp
 
; (38)

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control


p
-pressure in
chambers 10.
For a constant flight regime, altitude and airspeed (
const., const.HV

 ), which mean that
the air pressure and temperature before the engine’s compressor are constant


**
11
const., const.pT
, the term in equation (36) containing
*
1
p
becomes null.
3.3 System transfer function
Based on the above-presented mathematical model, one has built the block diagram with
transfer functions (see figure 5) and one also has obtained a simplified expression:

 
2
s1 1 s1
rpy rpy
y
Mc
p
n

py r
y
Mcpn Mci
pp
kk
k
kk
p
kk

  
   , (41)
where
rxl
kkk . 1s
M

k
c

+
n
py
k
pn
k
i

n
p
c
p
+
p
k

k

_

Fig. 5. System’s block diagram with transfer functions
So, one can define two transfer functions:
a.
with respect to the dosage valve’s lever angular displacement


sH

;
b.
with respect to the preset reference pressure
ci
p
, or to the sensor’s spring’s pre-
compression z,


s

yMcpn
p
r
py
r
py
yM cpn y M cpn
pp
k
kk
k
H
kk kk
kk kk
kk



  
 





   





10
rpy
cpn
p
kk
kk
k

. (45)
The first condition (43) is obviously, always realized, because both
y

and
M

are strictly
positive quantities, being time constant of the actuator, respectively of the engine.
The (40) and (41) conditions must be discussed.
The factor 1
cpn
kk is very important, because its value is the one who gives information
about the stability of the connection between the fuel pump and the engine’s shaft
(Stoicescu&Rotaru, 1999). There are two situation involving it:
a.
1
cpn
kk  , when the connection between the fuel pump and the engine shaft is a stable
controlled object;
b.
1

kk
k

and
10
rpy
M
p
kk
k






, which means that both other stability requests, (44) and (45), are
accomplished, that means that the system is a stable one for any situation.

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

315
If
1
cpn
kk 
, the factor
1
cpn
kk

M
cpn
kk
k
kk









, (46)
which offers a criterion for the time constant choice and establishes the boundaries of the
stability area (see figure 6.a).
Meanwhile, from the inequality (45) one can obtain a condition for the sensor’s elastic
membrane surface area’s choice, with respect to the drossels’ geometry (
,
An
dd) and quality


,
nA


, springs’ elastic constants


kk
k
l
S
kl k dp
 




. (47)
Another observation can be made, concerning the character of the stability, periodic or non-
periodic. If the characteristic equation’s discriminant is positive (real roots), than the
system’s stability is non-periodic type, otherwise (complex roots) the system’s stability is
periodic type. Consequently, the non-periodic stability condition is


2
11 4 0
rpy rpy
cpn y M yM
pp
kk kk
kk
kk


 

 







, (49)





222
2
12
2
prpy
p
pcpn rpy p
y
M
rpy p rpy
kkk kk k k kk
kkk kkk







, as well as vice-versa (until the stability conditions are
accomplished), assures the non periodic stability, but comparable values can move the
stability into the periodic domain; a very small
y

and a very big
M

are leading, for sure,
to instability.

UNSTABLE
SY STEM
STA B L E
SY STEM
NON-PERIODIC
STA BI L I TY







M
pyrppyr
ppyrpncp
y
kkkkkk
kkkkkkkkk

2
222
NON-PERIODIC
ST A B I L I TY
PERIODI C
STABILITY


M
pnc
p
pyr
y
kk
k
kk

1
1












p
cpn
k
p
tt
kk
k
kk













, (51)




1
cr
pcpnrpy
kkk

) during their stabilization.

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

317
The chosen RD-9 controller assures both stability and asymptotic non-periodic behavior for
the engine’s speed, but its using for another engine can produce some unexpected effects.

0
0.04
0.08
-0.04
-0.08
-0.12
-0.16
012
3
4
5
p
c
n
n
p
c
t [s]
0
0.05
0.10
-0.05

Q being
determined by the dosage valve’s opening.
As figure 8 shows, a rotation speed control system consists of four main parts: I-fuel pump with
plungers (4) and mobile plate (5); II-pump’s actuator with spring (22), piston (23) and rod (6);
III-differential pressure sensor with slide valve (17), preset bolt (20) and spring (18); IV-dosage
valve, with its slide valve (11), connected to the engine’s throttle through the rocking lever (13).

III
IV
n
I
3
4
5
6
7
8
9
10
11
12
0
0

y
FUEL TANK
i
p
0>z>0
0<x<0


ea
k
A
Q
B
Q
A
Q
S
T
O
P
i
Q
i
p
fuel
(to the combustor i njectors)
i
Q
i
p
p
Q
B
Q
c
p
0

opening (proportional to the (13) rocking lever’s angular displacement

).
4.1 Mathematical model and transfer function
The non-linear mathematical model consists of the motion equations for each above
described sub-system In order to bring it to an operable form, assuming the small
perturbations hypothesis, one has to apply the finite difference method, then to bring it to a
non-dimensional form and, finally, to apply the Laplace transformer (as described in 3.2).
Assuming, also, that the fuel is a non-compressible fluid, the inertial effects are very small,
as well as the viscous friction, the terms containing m,

and

are becoming null.
Consequently, the simplified mathematical model form shows as follows





s1
p
BA px
pp
kx

, (53)

BAAB
pp

ip
QQ k


 , (57)

pp
n
py
Qknk
y
. (58)
The model should be completed by the jet engine as controlled object equation


*
1
s1
MciHV
nkQ k p

  , (59)
where, for a constant flight regime, the term
*
1HV
k
p
becomes null.
The equations (53) to (59), after eliminating the intermediate arguments
,,,







  


.(60)

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

319
System’s transfer function is


sH

, with respect to the dosage valve’s rocking lever’s
position

. A transfer function with respect to the setting z,


s
z
H , is not relevant, because
the setting and adjustments are made during the pre-operational ground tests, not during
the engine’s current operation.

1
11
px py
pcpnM
AB pic Qx
kk
gkk
kk k





  





,


0
1
px py
cpn
AB
p
ic Qx
kk

before the step input, but the system is a static-one and it’s affected by a static error, so the
new value is, in this case, higher than the initial one, the error being 4.2%. The engine’s
speed has a different dynamic behavior, depending on the
cpn
kk particular value.

0 1 2 3 4 5 6
t [s]
p
c
p
i
0 1 2 3 4 5
0
t [s]
n
a)
b)
-0.05
-0.04
-0.03
-0.02
-0.01
0
0.01
0.02
0.03
0.04
0.05
0.01

n



Fig. 9. System’s quality (system time response for

-step input)
One has performed simulations for a VK-1-type single-spool jet engine, studying three of its
operating regimes: a) full acceleration (from idle to maximum, that means from

Aeronautics and Astronautics

320
max
0.4 n to
max
n ); b) intermediate acceleration (from
max
0.65 n

to
max
n ); c) cruise
acceleration (from
max
0.85 n

to
max
n ).

5
6
7
8
9
10
11
12
13
14
15
16
17
18
19
l
1
l
2
l
3
l
4



0>u>0
0>x>0
0<y<0
Q

i
p
e
p
p
20
l
6
l
5
VIII
21
22
23
24
VII
25
26
27
28
29
p
f
*
0<z<0
p
*
1
p
*

p
i
pp

to the throttle’s position (given by the 1
lever’s

-angle).
The fuel flow rate
i
Q , injected into the engine’s combustor, depends on the injector’s
diameter (drossel no. 15) and on the fuel pressure in its chamber
i
p
. The difference
p
i
pp
 ,
as well as
i
p
, are controlled by the level of the discharged fuel flow
S
Q through the
calibrated orifice 10, which diameter is given by the profiled needle 11 position; the profiled
needle is part of the actuator’s rod, positioned by the actuator’s piston 9 displacement.
The actuator has also a distributor with feedback link (the flap 13 with its nozzle or drossel
14, as well as the springs 12), in order to limit the profiled needle’s displacement speed.
Controller’s transducer has two pressure chambers 20 with elastic membranes 7, for each

p
kp kxk y

 
, (63)

22 2
p
pi RR Q p
p
k
p
k
p
kQ 
, (64)

y
RR ypp
y
k
p
k
p

, (65)

33ippy
p
k

 
. (69)
where the used annotations are

2
16
00
1
,
4
2
PT d
pi
d
k
pp





0
44
2
,
R
xx
p
kd



Aeronautics and Astronautics

322


2
7
7
00
1
,
4
2
RP
pR
d
k
pp






22
0
01 10
2
4tantan


,
11 11 0
0
1
() ,
2
zR s
R
kdzz
p




440
0
1
()
2
xR s
R
kdxx
p


 ,
0
0
u

 ,
0
0
i
di
d
p
k
p
 ,

0
1
0
RP p
p
RP xR zR R
kp
k
kkk
p


,
2
1
R
y
zz
Sl

0
Qi i
i
i
k
p
k
Q



0
2
0
PT i
p
RP PT
p
kp
k
kkp


,

0
2
0
RP R
R

k
kk
y


,

0
120
Pp
yp
rr
Sp
k
kk
y


,

0
3
0
PT p
p
PT si Qi i
kp
k
kkkp




. (70)
Furthermore, if the input signal u is considered as the reference signal forming parameter,
one can obtain the expression



ref
xd d d
xk
pp
  , (71)
where
re
f
d
p
is the reference differential pressure, given by
4
30
ref
rs
dr
md
kkl
pk
Slp



xd
k
y
R
p
p
p
p
p
p
p
p
p
3p
k

3y
k
y

2p
k
i
p
i
p

c
k
i

p
p
p
u
k

u

k
u
+
_
_
_
_
+
+
_
_
+
+
+
+
+
+
x
HV
k
*
1

1
p
n
d
p
y
n

Fig. 11. Block diagram with transfer functions

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

323
5.2 System quality
As figures 10 and 11 show, the basic controller has two inputs: a) throttle’s position - or
engine’s operating regime - (given by

-angle) and b) aircraft flight regime (altitude and
airspeed, given by the inlet inner pressure
*
1
p
). So, the system should operate in case of
disturbances affecting one or both of the input parameters


*
1
,
p

-0.01
-0.005
0
0.005
0.01
012345678910
t [s]
-0.0091
-0.0232
y(t)
n(t)
0.0087
p
d
(t)
-2
0
2
4
6
8
10
12
14
x 10
-3
012345678910
0,011
n(t)
y(t)

stabilizing at their new values with static errors, so the system is a static-one. However,
the static errors are acceptable, being fewer than 2.5% for each output parameter. The
differential pressure and engine’s speed static errors are negative, so in order to reach the
engine’s speed desired value, the throttle must be supplementary displaced (pushed).

Aeronautics and Astronautics

324
For immobile throttle and step input of
*
1
p
(flight regime), system’s behavior is similar (see
figure 12.b), but the static errors’ level is lower, being around 0.1% for
d
p
and for y , but
higher for
n (around 1.1%, which mean ten times than the others).
When both of the input parameters have step variations, the effects are overlapping, so
system’s behavior is the one in figure 13.a).
System’s stability is different, for different analyzed output parameters:
y
has a non-
periodic stability, no matter the situation is, but
d
p
and n have initial stabilization values
overriding. Meanwhile, curves in figures 12.a), 12.b) and 13.a) are showing that the engine
regime has a bigger influence than the flight regime above the controller’s behavior.

(t)
y(t)
0.0084
-0.06
-0.05
-0.04
-0.03
-0.02
-0.01
0
0.01
012345678910
t [s]
y(t)
0.00723
-
0.00362
p
d
(t)
-0,0538
n(t)
a)
b)

Fig. 13. Compared step response between a) stabile fuel pump-engine connection

1
cpn
kk 

As fig. 14 shows, the engine speed
n and the combustor temperature
*
3
T (see figure 14.b), as
well as the fuel differential pressure
d
p
and the pump discharge slide-valve displacement
y

(see figure 14.a) have periodic step responses and significant overrides (which means a few
short time periods of overspeed and overheat for each engine full acceleration time).
The above-described situation could be the consequence of a miscorrelation between the
fuel flow rate (given by the connection controller-pump) and the air flow rate (supplied by
the engine’s compressor), so the appropriate corrector should limit the fuel flow injection
with respect to the air flow supplying.

0 2 4 6 8 10 12 14 16 18 20
x 10
-3
-10
-8
-6
-4
-2
0
2
4
6

equipment (correctors), one for the flight regime and the other for the fuel-air flow rates
correlation.
The correctors have the active parts bounded to the 13-lever (hemi-spherical lid’s support of
the nozzle-flap actuator’s distributor). So, the 13-lever’s positioning equation should be
modified, according to the new pressure and forces distribution.
5.3.2 Barometric corrector
The barometric corrector (position VII in figure 10) consists of an aneroid (constant pressure)
capsule and an open capsule (supplied by a
*
1
p
- total pressure intake), bounded by a
common rod, connected to the 13-lever.

Aeronautics and Astronautics

326
The total pressure
*
1
p
(air’s total pressure after the inlet, in the front of the engine’s
compressor) is an appropriate flight regime estimator, having as definition formula


**
1
,
HHc
pp M





.
The new equation of the 13-lever becomes



2
*
5
12 1
2
2
dd
d
d
RR pp s r r H a
yy
l
Sp Sp m k k y S p p
tl
t

   
, (73)
where
a
p


.
5.3.3 Air flow-rate corrector
The air flow-rate corrector (position VIII in figure 10) consists of a pressure ratio transducer,
which compares the realized pressure ratio value for a current speed engine to the preset
value. The air flow-rate
a
Q is proportional to the total pressure difference
**
21
p
p

, as well as
to the engine’s compressor pressure ratio
*
*
2
*
1
c
p
p


. According to the compressor universal
characteristics, for a steady state engine regime, the air flow-rate depends on the pressure
ratio and on the engine’s speed



, from an intermediate compressor stage “f”, should be used, the intermediate pressure
ratio
*
*
*
1
f
f
p
p


being proportional to
*
c

. The intermediate stage is chosen in order to obtain

Aircraft Gas-Turbine Engine’s Control Based on the Fuel Injection Control

327
a convenient value of
*
f
p
, around
*
1
4
p

28
29
c
f
S
pp
S

, (75)
where
28 29
,SS
are 28 and 29-drossels’ effective area values. Consequently, the transducer
operates like a
*
c

-based corrector, correlating the necessary fuel flow-rate with the
compressor delivered air flow-rate. So, the corrector’s equations are:

**
22
()
p
pn
or
**
(),
ff
pp

2
dd
d
d
RR pp s r r mp c
yy
l
Sp Sp m k k y S p p
tl
t

    , (78)
where
mp
S is the transducer’s membrane surface area. After linearization and Laplace
transformer applying, its new non-dimensional form becomes





22 * *
011
s2 s1 ,
yy yRRypptfft
TT
y
k
p
k

*
0
*
28
29 0 26
mp f
f
r
fr
Sp
S
k
Sxk


 ,
 
**
010
66
1
1202 12102
,
mp f sf mp s
tf t
rr rrp
Spk Spk
ll
kk
kk

yy


 , (81)

Aeronautics and Astronautics

328
which should replace the (65)-equation in the mathematical model (equations (63) to (69)).
The new block diagram with transfer functions is depicted in figure 11.
5.3.4 System’s quality
System’s behavior was studied comparing the step responses of a basic controller and the
step response (same conditions) of a controller with correctors. Fig. 15.a presents the step
responses for a controller with barometric corrector, when the engine’s regime is kept
constant and the flight regime receives a step modifying. The differential pressure
d
p

becomes non-periodic, but its static error grows, from -0.1% to 0.77% and changes its sign.
The profiled needle position
y behavior is clearly periodic, with a significant override, more
pulsations and a much bigger static error (1.85%, than 0.2%). Engine’s most important
output parameter, the speed
n, presents the most significant changes: it becomes non-
periodic (or remains periodic but has a short time smaller override), its static error
decreases, from 1.1% to 0.21% and it becomes negative.
However, in spite of the above described output parameter behavior changes, the
barometric corrector has realized its purpose: to keep (nearly) constant the engine’s speed
when the throttle has the same position, even if the flight regime (flight altitude or/and
airspeed) significantly changes.

0
0.02
0.04
0.06
02468101214161820
T
3
(t)
*
n(t)
b)
with air flow rate corrector
without air flow rate corrector
-0.005
0
0.005
0.01
0.015
0.02
0.025
0.03
012345678910
t [s]
p
d
(t)
y(t)
-
0,0021
0.0077

Fuel injection is the most powerful mean to control an engine, particularly an aircraft jet
engine, the fuel flow rate being the most important input parameter of a control system.
Nowadays hydro-mechanical and/or electro-hydro-mechanical injection controllers are
designed and manufactured according to the fuel injection principles; they are
accomplishing the fuel flow rate control by controlling the injection pressure (or differential
pressure) or/and the dosage valve effective dimension.
Studied controllers, similar to some in use aircraft engine fuel controllers, even if they
operate properly at their design regime, flight regime’s modification, as well as transient
engine’s regimes, induce them significant errors; therefore, one can improve them by adding
properly some corrector systems (barometric and/or pneumatic), which gives more stability
an reliability for the whole system (engine-fuel pump-controller).
Both of above-presented correctors could be used for other fuel injection controllers and/or
engine speed controllers (for example for the controller with constant pressure chamber), if
one chooses an appropriate integration mode and appropriate design parameters.
7. References
Abraham, R. H. (1986). Complex dynamical systems, Aerial Press, ISBN #0-942344-27-8, Santa
Cruz, California, USA
Aron, I.; Tudosie, A. (2001). Jet Engine Exhaust Nozzle’s Automatic Control System,
Proceedings of the 17
th
International Symposium on Naval and Marine Education, pp. 36-
45, section III, Constanta, Romania, May 24-26, 2001
Jaw, L. C.; Mattingly, J. D. (2009).
Aircraft Engine Controls:Design System Analysis and Health
Monitoring,
Published by AIAA, ISBN-13: 978-1-60086-705-7, USA
Lungu, R.; Tudosie, A. (1997). Single Jet Engine Speed Control System Based on Fuel Flow
Rate Control,
Proceedings of the XXVII
th

Plasma-Assisted Ignition and Combustion
Andrey Starikovskiy
1
and Nickolay Aleksandrov
2
1
Princeton University
2
Moscow Institute of Physics and Technology
1
USA
2
Russia
1. Introduction
The history of application of thermally-equilibrium plasma for combustion control started
more than hundred years ago with IC engines and spark ignition systems. The same
principles still demonstrate high efficiency in different applications. Recently particular
interest appears in non-equilibrium plasma for ignition and combustion control
[Starikovskii, 2005; Starikovskaia, 2006]. The reason of the interest rise is new possibilities
for ignition and flame stabilization which are proposed by plasma-assisted approach. Over
the last decade, significant progress has been made in understanding the mechanisms of
plasma-chemistry interaction, energy redistribution in discharge plasma and non-
equilibrium initiation of combustion. Wide range of different fuels has been examined using
different types of discharges.
There are several mechanisms to affect a gas when using a discharge to initiate combustion
or stabilize a flame. There are two thermal mechanisms: 1) gas heating due to energy release
leads to an increase in the rates of chemical reactions; 2) inhomogeneous gas heating
generates flow perturbations and provokes turbulization and mixing. Non-thermal
mechanisms include 3) the effect of ionic wind (momentum transfer from electric field to the
gas due to space charge-electric field interaction); 4) the ion and electron drift in the electric

suitable for flight based on ram air compression. The lower boundary of this envelope is set
by thermal and structural limitations and is typically limited by a dynamic pressure of about
1 atm.
Relatively high dynamic pressure q is required, compared to a rocket, to provide adequate
static pressure in the combustor. The upper boundary is characterized as a region of low
combustion efficiency and narrow fuel/air ratio ranges thereby is restricted by combustion
stability in the engines (dynamic pressure limit is 0.25-0.5 atm).
The high Mach number (M > 15) edge of the envelope is a region of strong leading shock
waves, with strong dissociation and ionization of the gas in the shock layer. Here,
nonequilibrium flow can influence compression ramp flow, induce large leading-edge
heating rates, and reduce the efficiency of fuel injection and mixing and combustion. This
lead to dramatic performance decrease and puts another limitation on the possible flight
envelope for air-breathing hypersonic vehicles. On the contrary, at very low Mach numbers
(M < 3) the compression ratio due to flow deceleration is not enough for efficient ramjet
operation.
That is why in the low-speed regime (M = 0-3) the vehicle may utilize one of several possible
propulsion cycles such as a Turbine Based Combined Cycle (TBCC) with a bank of gas
turbine engines in the vehicle, or Rocket Based Combined Cycle (RBCC), with integrated
T
total
(K)
1
5
0
0
2
7
5
0
4

a
m
i
c

P
r
e
s
s
u
r
e
,

a
t
m
76,250
61,000
45,750
30,500
15,250
0
A
l
t
i
t
u

(Figure 2) [Andreadis, 2005]. As the vehicle continues to accelerate beyond Mach 7, the
combustion process is unable to separate the flow and the engine operates in scramjet mode