Ch6 Handout

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.


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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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.


Subsonic Inlets

This figure shows the locations in theengine intake where separation is most

likely to take place.


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Subsonic Inlets


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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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*


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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Supersonic Inlet

Shock-Boundary layer Interaction


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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Supersonic Inlet

External Deceleration


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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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 = = =




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20% is fed into the

primary zone of

which 12% passes

through swirling

vanes


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Afterburner and Ramjet combustors




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22

Exhaust Nozzle


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Exhaust Nozzle


Exhaust Nozzle


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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.


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

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Ch6 Handout

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