Aircraft propulsion combustor diffusor

Aeropropulsion
Unit
Combustor and Diffuser Design
2005 - 2010
International School of Engineering, Chulalongkorn University
Regular Program and International Double Degree Program, Kasetsart University
Assist. Prof. Anurak Atthasit, Ph.D.

Aeropropulsion
Unit
2
A. ATTHASIT
Kasetsart University
Topics
1.Combustion chamber design approach
2.Pressure loss in combustion chamber
3.Combustion chamber description
4.Combustor’s diffuser
5.Performance criteria

Aeropropulsion
Unit
3
A. ATTHASIT
Combustion Chamber : Problem
Compressor outlet velocity 170 m/s
Combustion chamber :
1.High velocity : impractical to attempt to burn fuel in air flowing at such high velocities
2.High velocity : high loss due to high friction
High inflow : impractical to attempt to burn fuel in air flowing at such high velocity

Aeropropulsion
Unit
Kasetsart University A. ATTHASIT 5
Pressure Drop in Burner
burner cold hot P P P
Combustor design requirements:
Minimizing the pressure drops

Aeropropulsion
Unit
Pressure Drop in Burner
Cold loss: sum of the losses arising in
the diffuser and the liner
burner ,cold 2
max
P
k M
P 2
 

Hot loss: occurs whenever heat is added
to a flowing gas
2 4
burner ,hot
3
T
P 0.5 U 1
T
 
 
   
 
Ref: Roffe, Gerald and Venkataramani (1978)

Aeropropulsion
Unit
Pressure Loss – Cold Loss
burner ,cold 2
max
P
k M
P 2
 

Ref: Roffe, Gerald and Venkataramani (1978)
Cold loss: the total pressure loss which results from sudden
expansions and contractions in flow are
k resistance coefficient (function of the system geometry)
γ specific heat ratio
Mmax the highest Mach number achieved in the contraction based on
the area ratio and mass flow

Aeropropulsion
Unit
Pressure Loss – Cold Loss
ref
max
M
M
1 B / 1000


Mref the reference Mach number
B the flameholder blockage (50-80%)
the ratio of the total blockage area
at the station where the base of the
flameholder is located to the
reference area of the combustor

Aeropropulsion
Unit
9
A. ATTHASIT
Combustion Chamber – Design Limitation
Flame front
Air-fuel ratio 18:1 – 24:1
Low efficiency

Aeropropulsion
Unit
10
A. ATTHASIT
Combustion Chamber : Solutions
?
- Burn it slowly - Increasing the reaction surface

Aeropropulsion
Unit
11
A. ATTHASIT
Combustion chamber - Anatomy
Vair inlet 170 m/s
Diffuser :
reduce 5 time
air speed
Recirculation zone : Primary zone: air-fuel ratio 18-24
Diluted zone:
burned product + fresh air

Aeropropulsion
Unit
13
A. ATTHASIT
Dilute Zone
Dilute zone:
Fresh air is mixing with the hot burned products to reduce their temperature to a value that is acceptable to the turbine

Aeropropulsion
Unit
14
A. ATTHASIT
Diffuser
The function of the diffuser is not only to reduce the velocity of the combustor inlet air, but also to recover a much of the dynamic pressure as possible, and to present the liner with a smooth and stable flow

Aeropropulsion
Unit
15
A. ATTHASIT
Diffuser – 2 types
Flare (or step) Diffuser
Dump Diffuser

Aeropropulsion
Unit
16
A. ATTHASIT
Influence of divergence angle on pressure loss
Divergence angle
Loss
Stall loss
Friction loss
Short diffuser:
High divergence angle
Stall losses arising from boundary layer separation
Long diffuser:
Low divergence angle
Pressure loss is high due to skin friction along the walls
Optimum: 6-12 deg.

Aeropropulsion
Unit
Diffuser Geometry
W
N
θ
L
1
L
AR 1 2 sin
W
  
R N
L
θ
2
1 1
L L
AR 1 2 sin sin
R R
 
 
   
 
Two-dimensional type
Conical-type

Aeropropulsion
Unit
Diffuser Geometry
W
N
θ
L
R N
L
θ
AR Area ratio, the primary function of the
diffuser achieving a prescribed
reduction in velocity
L/W or L/R non dimensional length, defines the
overall pressure gradient (the principal
factor in boundary layer development)
2θ the divergence angle

Aeropropulsion
Unit
19
A. ATTHASIT
Flow Regime in the Diffuser
1.No stall
2.Transitory stall: the eddies are formed + pulsating flow between core and boundary layer
3.Fully developed stall: the major portion of the diffuser is filled with a large triangular shaped recirculation region
4.Jet flow (occurs only at high angles of divergence)

Aeropropulsion
Unit
Performance Criteria
Obj: to understand the efficiency
parameters related to the diffuser
Mean velocity
m
u
 A

Dynamic pressure
2 u
q
2


Pressure loss :
(internal energy loss
+ redistribution of velocity between
inlet and outlet)
diff t1 t2 P  P  P
Area Ratio (AR) :
2 1 1
1 2 2
A u u
AR
A u u


  

Aeropropulsion
Unit
Static Pressure Rise in Diffuser
Several useful parameters for
expressing diffuser performance can
be derived from this equation
1 1 2 2 diff
2 1 1 2 diff
p q p q P
1
p p q 1 P
AR


   
 
     
 
Bernoulli:

Aeropropulsion
Unit
Pressure-Recovery Coefficient
  2 1
p
p p
C
q


Specific case: No losses (ideal pressure-recovery coefficient)
2 1 1 2 diff
1
p p q 1 P
AR

 
     
  No pressure loss
  2 1ideal
p 2
1
p p 1
C 1
q AR
  
    
 
(Dependent solely on area ratio)

Aeropropulsion
Unit
Overall Efficiency
p,measured
p,ideal
C
C
 
0.5   0.9
Overall efficiency varies depending on the geometry
and flow conditions

Aeropropulsion
Unit
Loss Coefficient
t1 t 2
1
P P
q



Where
t 2
t 2
2
P
P
m
 Constant mass flow
t1 t 2
1
P P
q



-Flare diffusers  λ~0.15
-Dump diffusers  λ~0.45

Aeropropulsion
Unit
Kinetic Energy Coefficient
With non uniform flows the kinetic energy flux is greater than it
would be for the same flow rate under uniform flow conditions
2
2
1
u udA
2
1
u m
2

 

α=1 completely uniform flow
α=2 flow on the point of separation

Aeropropulsion
Unit
Kinetic Energy Coefficient
Recall: under uniform flow condition α=1 completely uniform flow
1 1 2 2 diff p  q  p  q P
non uniform flow condition α≠1
1 1 1 2 2 2 diff p  q  p  q P

Aeropropulsion
Unit
Diffuser Performance
Parameters: Ch13P01
2
p,ideal 2
1
2 1
p
1 1
2 1
2
1 1 2
p,ideal p
1
C 1
AR
p p
C
q
p p
q
AR
C C







 
   
 




 
  
 
 
Proof these expressions :
Prove
• Obj: Able to
use the
fundamental
equation
under the
correct
assumptions
Analysis
• Obj:
Understand
the physical
meaning of
each
parameters
Calculation
• Obj: Able to
solve the
relations
under the
constraints of
corrected unit,
constant, …
etc.

Aeropropulsion
Unit
28
A. ATTHASIT
Prediction of the Diffuser Performance
Range of validity
-The flow is subsonic, but not necessary incompressible
-The inlet Reynolds number is greater than 2.5x104, so that problems of transition from laminar to turbulent flow are avoided
-The inlet velocity profile is symmetric
- Flow within the diffuser is essentially unstalled
- The diffuser itself is symmetrical and non turning

Aeropropulsion
Unit
Prediction of the Diffuser
Performance : Conical Diffuser
R N
L
θ
Performance chart for conical diffusers B1=0.02
(adapted from Sovran and Klomp 1967)
Maximum pressure recovery for
non-dimensional length at given
area ratio
The diffuser non-dimensional
length, producing the maximum
pressure recovery at a prescribed
area ratio

Aeropropulsion
Unit
Performance Prediction: two-dimensional
diffusers
W
N
θ
L
Overall effectiveness
Effectiveness is diminished by an
increase in inlet boundary layer
thickness
Small effect of boundary layer on
pressure recovery
Reneau et al (1967)
Domain of validation:
5<2θ<30 deg.
N/W1 : 1.5-25

Aeropropulsion
Unit
31
A. ATTHASIT
Design a conical diffuser to give maximum pressure recovery under a non- dimensional length N/R1 = 4.66 using the data given by Sovran and Klomp 1967
Diffuser Preliminary Design: Ch13P02
Prove
•Obj: Able to use the fundamental equation under the correct assumptions
Analysis
•Obj: Understand the physical meaning of each parameters
Calculation
•Obj: Able to solve the relations under the constraints of corrected unit, constant, … etc.

Aeropropulsion
Unit
32
A. ATTHASIT
From the graph, using log-linear scaling, the appropriate value of Cp is 0.6 and the corresponding value of AR is 2.13.
Cpi = 1-(1/2. 132) = 0.78.
Hence, η = 0.6/0.78 = 0.77. Transposing the expression given in the given figure, the included cone angle can be found: 2θ = 2tan-1{(AR0.5 - l)/(L/Rl)} = 11.26deg.
Diffuser Preliminary Design: Ch13P02 - Solution

Aeropropulsion
Unit
33
A. ATTHASIT
Design a conical diffuser to give maximum pressure recovery at a prescribed area ratio AR = 1.8
Prove
Analysis
Calculation

Aeropropulsion
Unit
34
A. ATTHASIT
Diffuser Preliminary Design: Ch13P02 - Solution
From the graph, Cp = 0.6 and N/R1 = 7.85 (using log-linear scaling). Thus, 2θ = 2tan-1{(AR0.5 - l)/(L/Rl)} 2θ = 2tar1-1{(1.80.5 - 1)/7.85) = 5deg. Cpi = 1 - (1/1.82 ) = 0.69 and η= 0.6/0.69 = 0.87
Do not forget showing an interpolation procedure

Aeropropulsion
Unit
35
A. ATTHASIT
An annular diffuser with an area ratio, AR = 2.0 is tested at low speed and the results obtained give the following data:
at entry, α1 = 1.059, B1 = 0.109
at exit, α 2 = 1.543, B2 = 0.364, Cp = 0.577
Determine the diffuser efficiency.
NB B1 and B2 are the fractions of the area blocked by the wall boundary layers
at inlet and exit (displacement thicknesses) and are included only to illustrate the
profound effect the diffusion process has on boundary layer thickening
Prove
Analysis
Calculation

Aeropropulsion
Unit
Conclusion
*
2
1
*
2
1
1
*
*
2
*
1
2( 1)
2
*
1
2
1
1
2
1
2
1
1
2
1
2
1
1
2
1
1
1 2
1
2
T
T
M
P
P
M
P
P
T
M
T
P
m AV AM
R T
M
A
A M











 










  
 
 
  
 
  
           
  
 
   
           
 
    
    
    
  
 
 
2
0
0 t
dA d du
A u
udu dP
dh dh udu
dP d dT
P T
a
P







  
 
  
 

P dP
T dT
d
A dA
u du
 





P
T
A
u

dx
2
dP
P 
See You
Next Class!

Aircraft propulsion combustor diffusor

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