“The Aerodynamic Challenges of Aero -engine Gas-Turbine ... · PDF fileLECTURE TITLE AND...

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17_9_13 LANCHESTER_LECTURE 1/47

“The Aerodynamic Challenges of Aero-engine Gas-Turbine Combustion Systems”

Prof. J. J. McGuirk Dept. of Aero. & Auto. Eng.,

Loughborough University, UK. 1. Intro. & background 2. Expt. facilities/techniques 3. CFD tools – turbulence

4. Compressor/combustor aerodynamic coupling 5. Combustor/turbine interaction 6. Fuel injectors – aero- & hydro-dynamics, aero-acoustics

7. A brief forward look 8. Summary/Conclusions/Acknowledgements

AERO-ENGINE COMBUSTION SYSTEM


Combustor RR Trent 700

The combustor accounts for ~1% of engine cost & weight.

A330 engine

1st of Trent family

EIS: 1995

57% market share

1000 units sold (Nov. 2012)

RR website: “lowest lifecycle fuel burn, lowest emissions and lowest noise levels of any A330 engine”

COMBUSTOR DESIGN REQUIREMENTS


High combustion efficiency (>99.5%) – complete fuel burn to liberate all chemical energy as heat (high and low power)

Good combustion stability - combustor operates over a wide range of AFR’s

Ease of ignition - both on ground and high altitude re-light Low total pressure loss – maximise overall efficiency Clean exhaust – Emissions must meet future regulations

(CO, UHCs, NOx, smoke) Good temperature traverse quality – overall and radial - to

ensure max. performance and turbine life Thermal and mechanical integrity – high combustor life,

influences wall cooling, thermo-acoustic properties Low weight/cost Aerodynamics - mixing, cooling, atomisation, acoustic processes

RICH BURN COMBUSTOR STYLE


Rich burn primary zone. stable down to low power

Rapid dilution/mixing to minimise stoichiometric zones

Lean dilution zone, hot enough for no smoke + extra air for traverse control

Pre-diffuser - dump diffuser design

RQL idea

ENVIRONMENTAL IMPACT


ACARE TARGETS Reducing environmental impact

LEAN BURN - TECHNOLOGY


40%

40%

20% 15%

15%

70%

Large multi-stream swirlers Staged, separate, rich pilot

(for low power, stability) and lean main (for low Nox)


LEAN BURN – THERMO-ACOUSTICS Heat release from flame can couple with combustion

system acoustics All flames (but particularly Lean Burn) are prone to this 1. Unsteady heat release causes acoustic wave propagation 2. Pressure drop across injector fluctuates 3. Fluctuations in AFR of mixture reaching flame 4. (Possible) reinforcement of unsteady heat release

Unsteady Heat Release

Air and fuel mass flow

oscillations

Pressure oscillations

Aerodynamic and aero-acoustic properties of injector and cooling flows crucial

EXPERIMENTAL FACILITIES - 1


Single or simplified flow elements



Single sector or multi-sector (airflow/liquid flow)



Fully annular + compressor (airflow)


EXPERIMENTAL FACILITIES - 4 Fully annular with upstream compressor (airflow)

Bespoke 1.5 stage rotor – engine representative OGV, pre-diffuser, dump diffuser, combustor geometry and flow split

Rotor inlet axial vel ~75m/s(M~0.2), flow coeff. VA /UB ~0.4 OGV Rechord > 2x105

(r - ri)/(ro - ri)

(P-P

b)/(P b

-pb)

0 0.2 0.4 0.6 0.8 1-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

LU ExperimentHP Rig 635/2

~~

~

Five-hole probe: time average velocity vector , pressure CO2 gas tracing Hot-wire: time-resolved velocity LDA, PDA, PIV, PLIF

INSTRUMENTATION


1.75mm


AERO-ACOUSTIC EXCITATION

Converts 105 pneumatic watts into 104 acoustic watts (165dB)

Frequencies up to 1200Hz Single plane waves or spectrum Bespoke design of exponential

horn 5m long/0.75 tonne (fatigue !)

1.7fEXPT ~ fENGINE (T-O)

REFRACTIVE INDEX MATCHING - 1



REFRACTIVE INDEX MATCHING - 2

CFD AND TURBULENCE MODELLING -1


CFD algorithms/codes Two basic approaches – density-based/pressure-based Main combustor flows are low Mach (<0.2) Except for thermo-acoustic oscillations !! Approach adopted is maintain pressure-based algorithm

popular in combustion - ‘mildly compressible’ via p/ρ link Turbulence modelling - again two approaches – RANS

statistical modelling/Large Eddy Simulation – former much used but not appropriate for aspects of combustor flows

PD exit


CFD AND TURBULENCE MODELLING -2

RANS

LES

RESOLVED

MODELLED

COMPRESSOR/COMBUSTOR COUPLING - 1


35%

35%

30%

15%

15%

70%

Change in combustor geometry can change compressor/combustor aerodynamic matching

Interaction is two-way: Upstream/downstream – 3D wakes/turbulence of OGV flow

help re-energise PD boundary layers, delay separation Downstream/upstream – dump gap (d/h~0.8) helps

stagnation pressure potential field on combustor cowl drive flow towards PD walls

Shift of injector flow (~30% to ~70%) alter design rules?



d/h = 0.8



IGV Inlet Rotor Exit Axial Velocity Profiles PD Exit Static

Pressure Profiles d/h = 0.8



minj = 50% minj = 70%

Injector Inlet Plane total pressure loss relative to rotor exit

Unwanted increase in non-uniformity of injector feed at high flow rate

d/h = 0.8


COMPRESSOR/COMBUSTOR COUPLING - 5 Induced azimuthal variations d/h = 0.8

Rotor Exit (Wall) Static Pressure

OGV Exit Centreline Total Pressure

Rotor-Inner Annulus Loss +20%



minj = 30-50-70%

d

hX4

Dump Gap Ratios Investigated: d/hX4 = 0.8, 1.2 and 1.6


Azimuthal Variation Rotor Exit Static Pressure


Design Profile

Forced Profile

minj = 70%

Rotor Exit Azimuthal Profiles

d/hx4 λ 0.8 0.262 1.2 0.265 1.6 0.290

Overall Loss Coeff.


COMPRESSOR/COMBUSTOR COUPLING - 8 Aerodynamically-coupled approach to compressor

OGV/combustor PD design – the Integrated OGV (IOGV)

Modified OGV lean/sweep = manipulated secondary flow = increased area ratio (for same length)

Conventional OGV/Pre-Diffuser L/h1 = 2.23, Area Ratio = 1.6

Integrated OGV/Pre-diffuser L/h1 = 2.23, Area Ratio = 1.8

30%

35%

35%



OGV Exit Velocity Contours/Vectors

Datum OGV Loss: λ = 0.12

IOGV OGV Loss: λ = 0.14

Datum IOGV

Rotor Exit Annulus Loss (IOGV relative to Datum) Inner: -10% Outer: -20%

RR have used IOGV concept in TXWB


COMBUSTOR/TURBINE INTERACTION -1

Annular rig - engine geometry extended to include NGVs Integer no of NGVs per burner, burner CO2 gas sampling Unchoked NGVs, emphasis on upstream effect of NGVs on

combustor flow and combustor turbulence exit conditions


COMBUSTOR/TURBINE INTERACTION -2

Plane D3 NGV row

Outer annulus

Inner annulus

Combustor outer liner

Combustor inner liner

Combustor exit flow

NGV outer bleed

NGV inner bleed

Turbine disc bleed

Probes

Plane D4

Outer

Inner

row

Improve effectiveness of NGV cooling air-reduce cool air/hot gas mixing Need proper scaling in

isothermal rig


COMBUSTOR/TURBINE INTERACTION - 3 Effectiveness of ambient, isothermal, gas Tracing for simulation of combustor exit traverse

Similarity – a consequence of high turbulent dilution mixing in rich burn – not very sensitive to density change/reaction

u’/U lint /c (c=NGV chord)

Hot-wire data at

plane D3 upstream

of NGV

High pressure combusting

Ambient pressure Isothermal

( )−AFR AFRAFR


COMBUSTOR/TURBINE INTERACTION - 4

Density Ratio Investigation

Exit plane

Heated air/CO2

Heated air/CO2

Inner/Outer nozzle cooling - separate flow manifolds to allow correct scaling to match engine Momentum flux ratio - (ρCUC

2/ρHUH2)

Heated air/CO2 mixture used to vary density ratio

(R-Ri) / (Ro-Ri)

Effe

ctiv

enes

s,η

0.2 0.4 0.6 0.8 10

0.05

0.1

0.15

0.2

0.25

0.3

0.35


COMBUSTOR/TURBINE INTERACTION - 5

(R - Ri) / (Ro - Ri)0 0.1 0.2 0.3 0.4 0.5

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

Datum RIDN Holes DR = 1Datum RIDN Holes DR = 1.15Datum RIDN Holes DR = 1.29LU Alternative RIDN Holes DR = 1LU Alternative RIDN Holes DR = 1.15LU Alternative RIDN Holes DR = 1.29E

ffect

iven

ess,

η

Alternative Design

Datum

Improved cooling effectiveness - lower metal temperatures RR have adopted this concept for next engine to enter service

FUEL INJECTORS - 1


Initial studies carried out on simplified but representative geometry to enable detailed comprehensive measurements

Geometry

Combined SPIV & PLIF

LES predicted steady state

FUEL INJECTORS - 2


LES videos

FUEL INJECTORS - 3


Expt. validation

FUEL INJECTORS - 4


Industrial complexity of injector nozzle

Medium mesh: 7.2M, 340KUse LES, get kSGS/k

Coarse mesh: 1.2M, 34K Use RANS, get L/Δ, Ret

Fine mesh away from walls:12.2M, 411K

Fine mesh near the walls: 12.4M, 1.1M

LES mesh design strategy

FUEL INJECTORS - 5


LES video Expt. video

(cavitation) RIM data

FUEL INJECTORS - 6


Lanchester (1908) - Fig. 86 ! PIV data

-0.8

-0.4

0

0.4

0.8

r/D

U/U

0 0.2 00-0.2

FUEL INJECTORS - 7


Injector aero-acoustics – initial simple (but relevant) problem to test method before applying to injector

NB (mildly compressible) URANS CFD - adequate as problem is superposition of plane wave acoustics and turbulent mean flow (?)

Acoustic damping by simple orifice

Absorption Coeff. L/D~0.5

FUEL INJECTORS - 8


Aerodynamic response quantified by Impedance – Z (complex) Orifice L/D? (not predictable by any simple acoustic model):

0.5

10.0

5.0

10.0

5.0

0.5

Resistance Reactance

FUEL INJECTORS - 9


Single swirl stream injector:

Excellent agreement between CFD and Expts. – can now use to identify ‘quiet’ injector properties

Resistance

Reactance

FUEL INJECTORS - 10


Most difficult - liquid primary atomisation (no detailed measurements) - resolve unsteady dynamics of air/liquid interface

LES using Combined Level Set and Volume of Fluid approach

Effect of turbulence

Two-phase flow:

FORWARD LOOK - 1


Cooled cooling air:


FORWARD LOOK - 2 Compressor/Injector Integration:

Velocity Contours Capture Streamtubes

- Swirl vane wakes

- TDC Deficit (pillar feed)

- BDC increase (PD exit hub-biased)

- Right/Left bias (residual swirl)

Offset

b MOS

a M

b MIS

47% IjUniformity Improvement: 44% 27%

bi

bo


Compressor/Injector Integration: FORWARD LOOK - 3

70% 70% 85%

0

30

60

90

120

150180210

240

270

300

330

0

0.2

0.4

0.6

0.8

1

1.2

Integration

- Experimentally evaluated - Design methodology validated

Injector flow rate 70%


FORWARD LOOK - 4 Pressure Gain Combustion ‘Const. Press.’ vs ‘Const. Vol.’

Theoretically possible: PG: 40 to 50%, SFC: -15 to -25% Much work on detonation devices - recent promising results from deflagration - pulsed combustor (Cam/RR) ~ 20% pressure gain. Challenge is integration Some success with unsteady ejector to couple to turbine Unsteady timed fuel injection

Inlet Exit

Noise and mechanical integrity? Compressor?

Source: Prof. R Miller (Cam. U) and RR

SUMMARY


Illustrations have been given of the important and challenging aspects of aerodynamics that underpin G-T combustion system technology

This is particularly the case when future lean burn combustors are

considered, and when integration with compressor/turbine systems is included – smart management of the component interfaces is key

The strategic use of isothermal atmospheric experimental facilities from

single component to fully-annular is an extremely cost-effective approach to improve our understanding of the wide range of aerodynamic challenges and identify future design concepts

The use of spatially-resolved, time-varying measurements in well-chosen

industrially relevant test cases is important to provide validation data and build confidence in improved (expensive!) computational methods

The success of the Large Eddy Simulation approach to turbulence modelling

in capturing the highly 3D, unsteady aerodynamic processes of high swirl should lead to substantially increased future use fuel injectors.


ACKNOWLEDGEMENTS Many colleagues/organisations contributed to the work presented:

Loughborough UTC colleagues, particularly: o Jon Carrotte, Adrian Spencer, Paul Denman, Andrew Garmory,

Gary Page, Mark Brend, Ashley Barker, Duncan Walker, Mehriar Dianat, Feng, Xiao, Chris Ford, Jailin Li

Rolls-Royce colleagues, particularly: o John Moran, Steve Harding, Ken Young, Mark Jefferies, Barani

Gunasekeran, David Dunham, Jochen Rupp

Other University colleagues, particularly: o Rob Miller (Cambridge)

Funding Agencies, particularly:

o RR, EPSRC, EU, TSB

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