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8/13/2019 Ch6 Handout
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8/13/2019 Ch6 Handout
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3Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Subsonic Inlets
An jet engine must be provided with an air intake and a ducting system.
For a turbojet engine, the airflow entering the compressor should have a Mach
# between 0.4 to 0.7.
This means that if the turbojet installed in an aircraft flying at Mach # of 2, the
air intake should be designed in such away that you get a Mach # of 0.4 to 0.7
at the inlet of the compressor.
In this case the inlet of the engine will act like a diffuser.
When designing the inlet of a jet engine, it is important that the stagnation
pressure loss is small.
4Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Subsonic Inlets
Flow pattern: the flow pattern at the
inlet of an engine depends on
flight velocity: high speed flight
and low speed flight.
1. High speed flight (Example:
Cruise):
The intake mass flow rate
required by the engine is low.
This is accompanied by
external flow deceleration at
the inlet.
This requires less internal
pressure rise (p2-p1) and
hence less severe loading of
boundary layer.
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5Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Subsonic Inlets
2. Low speed flight (Example: take-
off):
The intake mass flow rate
required by the engine is
high.
This is accompanied by
external flow acceleration at
the inlet.
The internal pressure rise
(p2-p1) can be very large
which will cause boundary
layer separation and hence
a diffuser stall.
6Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Subsonic Inlets
This figure shows the locations in theengine intake where separation is most
likely to take place.
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7Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Subsonic Inlets
8Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
In supersonic flow, it is important to design the inlet so that the Mach #
entering the compressor is subsonic.
Therefore, the flow should be decelerated from supersonic to subsonic.
This can be don by either a normal shock wave or a couple of oblique shock
waves.
However, the loss across a normal shock wave is very large.
A couple of oblique shock waves would be better. The loss across oblique
shock wave is less than normal shock.
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9Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
When designing a
supersonic inlet, we need to
consider different operation
conditions.
Lets consider the
acceleration of a fixed
geometry convergent-
divergent nozzle (CDN)
a. Aa is determined by the flow
downstream the inlet.
b. Aa
is determined by the flow
at the throat of the CDN.
in this case, At = A*
10Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
c. M =1, a weak shock appear in
front of the CDN
d. Increasing the Mach number will
yield a bow wave.
once the shock wave is
established, the flow entering
the inlet is no longer isentropic.
Therefore, the geometry of the
Inlet of the CDN should be
changed to prevent the
formation of shock wave.
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11Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
Shock-Boundary layer Interaction
12Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
External Deceleration
022
01
033
02
044
03
2.26; 0.895
1.65; 0.945
0.67; 0.870
pM
p
pM
p
pM
p
= =
= =
= =
Normal
shock
waveoblique
shock
waves
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13Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Supersonic Inlet
External Deceleration
14Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
Recall that the
fuel-to-air ratio
The stoichiometricfuel-to-air ratio
(fstoich) can be
higher than actual
fuel-to-air ratio (f).
fstoich can be
calculation from
reaction of fuel
and air.
04 03
04 0304 03
04 03
( )
( )for
1500 K, 600K, 45,000kJ/kg, 0.02
a f a f R
p
f a
R R
R
m m h m h m Q
c h hh hm m f
Q Q
T T Q f
+ = +
= =
= = = =
9.525 kmole of air, = 28.96 kg/kmol 275.844 kg.
1 kmole of fuel, = 16.04 kg/kmol 16.04 kg.
16.040.0581
551.69
a
f
f
stoich
a
m
m
mf
m
=
=
= = =
M
M
4 2 2 2 2 2
air
CH +9.525* (0.21O + 0.79N ) CO +2 H O + 7.52 N
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15Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
The equivalence ratio is then
If the fuel-to-air ratio is similar to the stoichiometric fuel-to-air ratio the turbine
inlet temperature will be very high.
So we need to keep the fuel-to-air ratio as small as possible to prevent
excessive temperature in the turbine.
0.020.34
0.0582stoich
f
f = = =
16Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
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17Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
18Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
20% is fed into the
primary zone of
which 12% passes
through swirling
vanes
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19Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
Afterburner and Ramjet combustors
20Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
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21Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Gas Turbine Combustors
22
Exhaust Nozzle
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23Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Exhaust Nozzle
24Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Exhaust Nozzle
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25Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Sample Problems
1. A ramjet engine is being designed for flight Mach number 4.5 at an altitude
where the ambient pressure ant temperature is 9 kPa and 220 K. The Mach
number of the flow at the entrance of the burner is 0.3 and the burner has a
constant cross-sectional area. The combustion may be represented
approximately as heating of a perfect gas with constant specific heat ratio. The
stagnation temperature at the burner exit is 2600 K. Neglecting frictional effects
in the burner and considering the flow to be one-dimensional throughout,
estimate the Mach number of the gas leaving the burner. Determine also the
static and stagnation pressure loss in the burner due to heating (the ratio of
outlet and inlet pressure). Assume =1.4.
26Flight Propulsion I: Aerothermodynamics of inlets, combustor and Nozzles
Sample Problems
2. Consider a convergent circular nozzle with an inlet area of 0.45 m2 and an inlet
stagnation pressure and temperature of p0 = 300 kPa and T0 = 1400 K. The
mass flow rate throughout the nozzle is 100 kg/s and the stagnation pressure
at the exit of the nozzle is 2% lower than at the entrance of the nozzle. If the
nozzle flow is convergent and chocked and the specific heat ratio is 1.36, find
the following at the exit of the nozzle:
(a) The exit velocity(b) The exit pressure.
(c) The exit area and diameter.
(d) The Mach number at the entrance of the nozzle