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Transcript of ROMAGNOSI English presentation
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SAPIENZA UNIVERSITY OF ROME
SCHOOL OF AEROSPACE ENGINEERING
MASTERS DEGREE IN ASTRONAUTICAL ENGINEERING
LES OF COMBUSTION IN SUPERSONIC REGIME
FOR SCRJ APPLICATIONS
SUPERVISOR STUDENT
Prof. Claudio Bruno Luigi Romagnosi
ASSISTANT SUPERVISOR
Ph.D Antonella Ingenito
Ph.D Donato Cecere
Academic Year 2009/2010
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Goals of thesis
2
The analysis of mechanisms of vorticity and turbolent production
in the field with the ultimate goal to optimize the mixing and
anchor the supersonic flame
Validation of results using measurments from the HyShot project
[Rif. Report on the Hyshot Scramjet Experiments in the T4
Shock Tunnel, M. Frost, A. Paull, H. Alesi]
X-51 A Waverider New concept space launcher
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3
Introduction
- Ramjet Scramjet
- How SCRJ model engine works
- HyShot scramjet program
Numerical approach
-Mathematical model and simulation set-up
- Closure models (SGS / EDC)
- Numerical scheme (Weno35)
Simulation results
- Description of the fluid dynamic field
- Study of the vorticity production and diffusion terms
- Combustion analysis
Conclusions and future developments
Contents
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SCRAMJET = Supersonic Combustion RAMJET
RAMJETis the evolution of the turbojet which, based on the idea of
Ren Lorin (1913), have no rotating parts. The absence of
compressor and turbine allows higher temperature in the combustion
chamber.
RAMJET SCRAMJET
RAMJET limits: C.C works in the subsonic conditions sharp
slowdown of the flow in the air intake high
temperature in the C.C limit on the maximumflight speed (M 5)
Solution: keep a supersonic flow in the combustion chamber (SCRJ)
Why studying SCRAMJET?
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How a SCRJ model engine works
5
Future: advantages:
High flight speed (M=6-12) No need for carrying oxidizer on board:
SCRJ uses air (for new concept launcher)
Drawbacks:
Must be accelerated up to M=6 Low residence time in c.c.(10 -3 10 -4 s) mixing is critical
Air intake
Combustion
chamber
Thrust
plate
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HyShot scramjet program
6
HyShot is a research project developed at the University of Queensland
Centre for Hypersonics (UQ) in order to demonstrate the feasibility of
supersonic combustion via flight tests (jointly with US and UK)
1st stage (Terrier) tburnout = 6.4 s
V = 4000 km/h
h = 3.7 km
2nd stage (Orion) tburnout = 27 s
V = 8300 km/h
h = 56 km
Test Fuel: H2tinjection = 6 s
h = 35 - 23 km
M = 7.6 7.4
Trajectory data:
Apogee: h= 314 km
Mission profile:
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Simulation of HyShot combustion chamber
7
= 0.426 Air Hydrogen
Pressure [Pa] 82110 307340
Mach 2.79 1
Density [kg/m3] 0.2358 0.3020
Temperature [K] 1229 250
Sound speed [m/s] 682.9 1204.4
Flow speed [m/s] 1905.291 1204.4
Simulation Data from UQ ground testing in the T4 SWT (h = 28 km ; AOA = 0):
305 mm x 100 mm 300 mm x 75 mm x 9.8 mm 200 mm x 75 mm
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Mathematical model and simulation set-up
8
Solver: Explicit and compressible
Method: Finite difference (placed variables)
Numerical scheme:Runge-Kutta 3rd order(time integration)
Hybrid: Finite differences 4th order-WENO35 (spatial integration)
SGS Model: Fractal
Riemann problem solver: HLLC/HLLE
Boundary conditions: NSCBC (Navier-Stokes Characteristic
Boundary Condition)
Kinetic scheme: 9 involved species and 37 chemical reactions
Reactive N-S:
Species transport
equations:
Eqn of state:
No.nod
es=
5
010
6
(448
x12
8x878)
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SGS (SubGrid Scale) closure models
9
Fractal nature of turbulence:
Hp: large Re
inertial range below
(eddy viscosity) with
Combustion model (EDC):
fine structures
V* = *V
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WENO35 numerical scheme
WENO (Weighted Essentially Non-Oscillatory) is the evolution of a scheme introduced for the
first time in 1987, developed by Harten, Osher, Engquist and Chakravarthy. WENO35 has third
order accuracy where the variables are discontinuous, and fifth order where smooth.
(candidate stencils for the reconstruction)
Case:r = 3 (5 cells) Accuracy: 2r-1 (smooth)
r (not smooth)Lagrange polynomials:
with
If the solution is
smooth in all Sk:
with
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WENO35 validation
11
PROGRESSIVE WAVE
REGRESSIVE WAVE
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Simu
lat i
on re
sults
(1/2)
H2 expands and
(vorticity generated
by baroclinic effect)
12
900 m/s
Mach disk
Barrel shock
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Simulation results (2/2)
M=2.40.6
T=250310 K
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Study of (vorticity)
Vorticity transport equation:
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Baroclinic term (1/3)
It is the only true source term of vorticity (as is
not a function of )
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Bar
oclin
icterm(2/ 3
)
16
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Bar
oclin
icterm(3/ 3
)
17
1
2
3
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Vortex Stretching (1/2)
The vortex stretching promotes the turbulence energy cascade through the
combined effect of stretching and tilting:
Rigid rotation does
not contribute to
vortex stretching
For example, to simplify matters:Incompressible fluid div(u)=0
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Vortex
Stre tc
hing( 2/
2)
UZ
= 200 - 1800
m/s
19
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Compressibility term (1/2)
Compressibility plays a dual role:
I. Reduces molecular mean free path shortens chemical time
II. Increases molecular collisions lower species interdiffusion(important for diffusion flames)
Mean free path:
Reaction rate [kg/m3s]:
k = ATb e
EA
/RT
(Arrheniuss kinetic theory)
Kelvins Theorem:
A = cost
L
L
(Incompressible fluid)
(Compressible fluid)
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Compre
ssib
ilit
yter m
(2/2)
21
div(u)>0
div(u)
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Viscous terms (1/3)Viscous terms are f(), diffuse vorticity and create small-scale vortices close
to the wall.
Dimensionaless form of vorticity equation:
with
NB: IfRe 1 then VISCOUS FORCES INERTIAL FORCES
Re 1 temperature rise
u flow slows down close to the wall
chemical reactionswall friction
Linked to thesecond derivativesof the vorticity. Itproduces vorticity
in opposition to the
vortex stretching
Lighter particles aresubjected to greaterdecelerations due to
viscous stress. Itproduces vorticity in
opposition to the
baroclinic term
Vortices directed in a generaldirection are redirected along
a definite direction whensubjected to viscous gradients
in the other two directions
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Viscous
ter
ms(2/3)
23
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Viscous
ter
ms(3/3)
24
COMPETITION BETWEEN
MASTER-SLAVE VORTICES
VS
Boundary layer separation at z = 53 mm
caused by p=8000 Pa in ~1 mm
V i i d Mi i (1/3)
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Vorticity and Mixing (1/3)
Vt = 1000 m/s
d = 2 mm
= 10-5 Pas
= 0.3 kg/m3
Re = 60000
= 500000 rad/s
K = LRe-3/4 0.5 m
t = TRe-1/2 50 ns tm
DIFFUSION
FLAME???
NOTE: NO KOLMOGOROV BUT FM (COMPRESSIBILITY)
= 105 106 Hz
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Vortic
ityan dM
ixin
g(2/3)
26
H l (T 250 k)
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Vortic
ityan dM
ixin
g(3/3)
= 80000 300000 rounds per second
H2 core very cool (T=250 k)
heating and consumption
from the outside
Competition between master slave vortices
instability of flame surface in favor of mixing
Redistribution of H2 along the
walls (tilting of spanwise
vorticity) increase in heat
transfer surface air/wall-H2
27
M i h i l i
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YH2 0.2 %
YOH 1.5 %
YH2O 10 %
Main chemical species
Si l ti lid ti
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Simulation: validationno. 16 pressure transducers
spaced 13 mm apart. The
first is located 9 cm
downstream of thecombustor chamber
entrance.
THRUSTPLATE
AIR INTAKE
COMBUSTION
CHAMBER
C l i
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Conclusions
The LES simulation of the HyShot II combustion chamber highlights some interesting
aspects:
this simulation predicts complete combustion in supersonic regime (flame anchors
already 2 cm upstream of the injectors)
crossflow injection allows rapid fuel-oxidant mixing; the baroclinic effect caused
by the expansion of the H2 jet produces high energy vortical structures
the baroclinic contribution is of the same order of magnitude of the vortex
stretching and compressibility terms (1010 rad/s2).
the hydrogen low density contributes to the production of vorticity (B is inverselyproportional to the square of )
combustion efficiency is very high (only 0.2% of the total mass at the
combustion chamber exit is H2)
F t d l t
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How much fuel injected affects vorticity production (for example,
kerosene: RP-1 800 kg/m
3
vs H2 0.09 kg/m
3
)
What is the thrust contribution by fuel momentum (for example,
vary the angle and the injection pressure)
How much the injector geometry affects the mixing (fluid jet
destabilization, injecting from slits)
What is the increase of entropy in different configurations (search
for the optimum set-up that gives minimum S). This simulation shows a
S of about 37/mol K through the combustion chamber
Future developments
arget: Looking for the right balance between mixing and thrust produced