23 - Air Breathing Engines - Combustors

8/18/2019 23 - Air Breathing Engines - Combustors

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Combustors

• In the combustor, the working fluid is heated (increase in stagnation temperature) by

chemical reactions, the stagnation temperature rise across the combustor depends on

the fuel/air ratio (f real) and the combustion efficiency (b)

• Liquid fuel (Jet‐A, JP‐8) is atomized and burned with air to form hot gases, the above

process is carried out in a combustor which facilitates stabilization of reactions, with

minimal stagnation pressure loss (~5% for gas turbine combustor)

• Gas turbine combustors are designed compact to reduce its weight and size and

typically have high heat release intensity (20‐50 MW/m3‐atm), with residence time for

gases ~10 ms, hence within a very short time, the fuel should atomize, vaporize, mixand burn with the air

• High combustion efficiency is expected (>99%) to minimize fuel wastage as well as to

reduce pollutant emissions (unburned hydrocarbon (UHC), carbon‐monoxide (CO))

• For gas turbine combustors, reduction in pollutant emissions (oxides of Nitorgen (NOx),

CO, UHC, soot) has gained interest and is one of the factors for combustor design due

to enforced emission norms for civilian aircrafts

Ramjet or afterburner

Gas turbine combustor


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Fundamental Stagnation Pressure Loss

• Constant area (A) duct, steady, 1D frictionless flow, assume calorically perfect gas

and constant properties of air (, cp)

• No flow enters or leaves from the combustor sides, consider negligible fuel mass

flow rate as compared to air mass flow rate so that the flow rate is assumed to besame at inlet and outlet

• u1, T1, P1 (T01, P01 can be calculated) and T02 and the inputs, knowledge of T02 and

T01 gives the heat addition rate

•

We want to find the stagnation pressure loss: Pb = (P01‐P02)/P01

Mass balance: 0

∀ ·

0, steady

0

Reactants Products

x

CS

CV

21

T1, P1 T2, P2

u1 u2

T02, P02T01, P01

Frictionless

A

Rayleigh flow:


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Fundamental Stagnation Pressure Loss

Linear momentum relation:

∑

∀ ·

0, steady

Energy conservation:

State and stagnation relations:

If T01 and T02 are inputs then f real can be calculated

1

1


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Example

• Given: u1 = 128.5 m/s, T01 = 486.81 K, P01 = 3.144 bar and T02 = 1200 K

• cp = 1148 J/kg‐K, = 1.333, R = 287 J/kg‐K

• Find: fundamental stagnation pressure loss due to heat addition: Pb = (P01‐P02)/P01

486.81

.

479.62 (M1 = 0.3)

.

. 2.152

.

..

.. 2.962

2.152 128.5 276.53

276.53 287 79364.11

Stagnation properties:

Density:

Mass conservation:

79364.11


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Example


2.962 10

276.53

128.5

2.962 10 276.53 35534.11

331734.11 276.53

331734.11 276.53

Stagnation temperature relation:

1

1200

1200

Substituting in mass conservation:

79364.11

331734.11 276.53 79364.11 1200


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Example

331734.11 276.53 95236932 34.57

241.96 331734.11 95236932 0

1371.37 393606 0

409 / (M2 = 0.62)

1200

1127.14

331734.11 276.53 409 2.186

2.186 .

.. 2.809

..

. 0.107 10.7%

(note: the other root will give supersonic speed neglected)

for: u1 = 128.5 m/s (M1 = 0.3)


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Example (lower inlet speed)

• note: significantly lower loss in the stagnation pressure if heat is added at lower inlet speeds

for: u1 = 43.12 m/s

486 (M1 = 0.1)

3.123 2.239

108.5 /

1194.87 (M2 = 0.16) 3.060

inputs: T01 = 486.81 K, P01 = 3.144 bar, T02 = 1200 K

3.113

0.98% for: u1 = 43.12 m/s (M1 = 0.1)


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Stagnation Pressure Loss

• Combustor stagnation pressure loss is due to (a)

friction and aerodynamics and (b) heat addition,

generally, frictional losses are ~ 2.5‐5% and heat

addition losses are ~ 0.5‐1%• Velocities at the exit of compressor are high

(~100‐150 m/s), hence, the flow needs to be

decelerated (~20‐30 m/s) before feeding it to

the combustor this is done in a diffuser

mounted upstream of the combustor Fig. A.7 Stagnation pressure loss in Rayleigh flow(Gas Turbine Theory by HIH Sarvanamuttoo, GFC

Rogers, H Cohen, Dorling Kindersley, New Delhi, 2009)

1

1

1

⋯

expanding in Taylor series:

1

⋯

1

⋯

⋯Higher order terms negligible for small M (


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Incompressible (M < 0.3)


Stagnation pressure, incompressible flow:

1

Mass conservation:

1 ≅

1 ≅

1inlet dynamic

pressure

≅

1

For M < 0.3

.

.

.

.

1.003

.

.

1.021

example:

example:


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Incompressible (M < 0.3)

≅

1

For u1 = 128.5 m/s

≅ .

2.152 128.5 .

1 8.3%

For u1 = 43.12 m/s

≅

.

2.152 43.12

. 1 0.93%

in this case: M1 = 0.3 and M2 = 0.62

in this case: M1 = 0.1 and M2 = 0.16

≅

1

. 1 1.47 (typical value: 1‐2)

loss in stagnation pressure due to heat addition is approximately 1.47 times the inlet

dynamic head for the example problem (it is typically 1‐2 times the inlet dynamic head)

~ 20

loss due to

heat addition:

loss due to

friction and

aerodynamics:

typically frictional losses are dominant as compared to heat addition losses


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Gas Turbine Combustor

Fig. 1.8 Derivation of conventional combustor configuration.

(Gas Turbine Combustion by Arthur H. Lefebvre, Second Edition, Taylor and Francis, New York, 1999)

Simplest possible form of combustor – straight walled duct

connecting the compressor and turbine, however, due to

high compressor outlet velocities (~170 m/s) fundamental

pressure losses would be very high

To reduce this pressure loss, a diffuser is used to lower the

velocity by a factor of about 5, still the air speed is high as

compared to flame speed (~1 m/s)

To provide a low velocity region combustion is sustained by

recirculatory flow of burned products (e.g. by use of swirler

or wake of a bluff body). The alloys used for combustor

cannot sustain high flame temperatures

Cooling arrangement (using film or back side cooling) is

facilitated by a part of inlet air. The combustor outlet

temperature is tailored by using dilution air to acceptable

limit of turbine blades


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Gas Turbine Combustor (air flow distribution)

Fig. 6.31 Air flow distribution in gas turbine combustor.

(Mechanics and Thermodynamics of Propulsion by Philip Hill and Carl Peterson, Second Edition,

Dorling Kindersley India Pvt. Ltd., Noida, 2010)

• About 12% of inlet air passes through swirling vanes that surround the fuel jet, inprimary zone, fuel/air ratio is nearly stoichiometric and is fed by other air flows

(8% and 20%)

• Swirling flow creates low static pressure region near the centerline and results in

recirculation of hot product gases and also facilitates low velocity region to

stabilize combustion

• In dilution zone (~20% of inlet air) the temperature of gases is reduced to a level

acceptable to turbine, uniform temperature at the combustor exit is desired to

minimize thermal stresses in the turbine blades

• Cooling of the combustor liner is facilitated by film cooling from a part of theinlet air (~40%)

23 - Air Breathing Engines - Combustors

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