piotr doerffer transonic flow control by streamwise vortices. research offer

114 June 201114 June 2011 – – Session III (P2)Session III (P2)

General Meeting of AERONET General Meeting of AERONET

Polish Academy of SciencesInstitute of Fluid Flow Machinery

Piotr Doerffer

Transonic flow control by streamwise vortices. Research offer

Cooperation between science and aerospace industry

1) Start-up of research – EPFL first investigations and patent applic.

2) Research at Gdansk

3) Inclusion of AJVG into AITEB2 project (coord. RRD), patent applic.

4) Inclusion of AJVG into UFAST project (coord. IMP PAN)

5) Further research in the FACTOR project

6) New concept of RVG, patent applic., research projects

2001/3

2003/4

2005/9

2005/9

2010/13

2011/14




One of very important flow cases

In transonic compressor cascades shock waves are formed

These shock waves interact with boundary layers and cause: -shock induced separation -unsteady effects of shock-boundary layer interaction

To limit these negative effects different flow control methods were tested, mainly for external aerodynamics




Reaearch idea from MIT ALSTOM Project for EPFL Lausanne

Transonic compressor blades are:

- very thin

- strong spanwise variation




Flow control concept – Stream-Wise Vortices (SV)

Methods of SV Generation

Vane Vortex Generators Air Jet Vortex Generators




Test Model and Measurement Techniques

pressure measurement through wall taps

PSP – pressure sensitive paint schlieren visualisation of shock

system oil visualisation

EFMC 2003, Toulouse,

Shock Wave – Boundary Layer Interaction Control by Stream-Wise Vortices

Piotr Doerffer – IMP PAN, Gdansk, Poland

Albin Bölcs, Klaus Hubrich - EPFL, Lausanne, Switzerland




20º 20º 20º 20º 20º

60° 60° 60° 60° 60° 60°

= 1 mm = 1 mm = 1 mm = 1 mm = 1 mm = 1 mm

= 0.4 mm = 0.4 mm

ALSTOM Patent

The Patent Application bears following data:

Our. Ref.: B03/172-0 DETitle of invention: Verfahren zur Verbesserung der Strömungsverhältnisse in einem Axialkompressor sowie Axialkompressor zur Durchführung des Verfahrens

Application No.: 103 55 108.5Filing Date: 24.11.2003

Legal Owner/Applicant: ALSTOM Technology Ltd,CH-5400 Baden




Own research at IMP PAN

0 540 X [mm]

Pivoted wall

Flat wall

Supply cavity

shock wave

Plate withthe AJVG

Location of AJVG

Traverse upstream the shock, 25 mm

Traverses downstream the shock, 30 and 55 mm

Measurement nozzle




Experimental investigations

Plate with the AJVG

VG holesStatic pressure holes

Main measurement:• Static pressure along the wall• Boundary layers in 3 traverses• Schlieren pictures• Oil visualisation• Oscillations of the shock wave

Mach numbers: 1.25, 1.35, 1.45, 1.55




Boundary layers 30 mm downstream

u/u_delta = F(y/y_delta), M=1.35

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

0 0.2 0.4 0.6 0.8 1

u/u_delta

y/y_

delt

a

BL for No JetsBL for Jets

u/u_delta = F(y/y_delta), M=1.45

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

0 0.2 0.4 0.6 0.8 1

u/u_delta

y/y_

delt

a

BL for No Jets

BL forJets




Shock wave oscillations

Power spectrum, M=1.35, Rear Shock

0.00

0.05

0.10

0.15

0.20

0.25

0 10 20 30 40 50 60

f [Hz]

Am

plit

ud

e

No JetsJets

RMS Rear shock

M No Jets Jets

1.35 0.95473 0.79997

1.45 1.13944 1.02838

1.55 2.18671 1.97621

Main shock

Rear shock




AITEB-2

Coordinator RRD




pitch/chord = 1.11 pitch/chord = 1.13

Increased blade load – increased pitchGeometry: IGG Mesh:

Chord: 120 mm 208,288 cells (y+ ~1)Axial Chord: 76.1 mm Numerical scheme:Central Difference Scheme: SPARC, FINE2nd Order Upwind: FLUENTTurbulence model: Spalart-Almaras

Boundary conditions for pitch/chord =1.13:Inlet: Outlet:Total pressure 27983 Pa Static pressure 15840 Pa Total temperature 303.3 KInlet angle 48.6 degViscosity ratio 10




Final adjustment of the shock location and smoothing of the blade shape




Reference case experimental results

Maximum Mach number reached

Obtained reduction of M upstream of the shock

Smaller -foot in experiment

Same shock location




Films with reference shock and with cooling and AJVG

Stabilisation of the shock wave




Cooling + AJVG

oil visualisation




y= f(PP/P0), X=128, Theta=45 alfa=90

02468

101214161820

0.45 0.55 0.65 0.75 0.85 0.95

PP/P0

y [

mm

]

reference Shock X=114Cooling on VG off Shock X=114Cooling on VG on Shock X=114

y= f(PP/P0), X=171, Theta=45 alfa=90

02468

101214161820

0.45 0.55 0.65 0.75 0.85 0.95

PP/P0

y [

mm

]

reference Shock X=114Cooling on VG off Shock X=114Cooling on VG on Shock X=114

Stagnation pressure

Application of cooling increases b.l. thickness considerablyAJVG decreases the effect

X = 128 X = 171




Objective function: maximum of vorticity X-component at the section 50mm downstream of the jet

Variables: jet skew angles „alfa” jet pitch angles „theta”

Optimisation of AJVG parameters




The invention submission EM 70518 by Rolls-Royce

dated 04th December 2008,

Title: “Method and apparatus to passively control flow separation on turbine airfoils by means of streamwise vortices generated by air jets”




Project starting date: 1st of December, 2005

Project duration: 3,5 years

Unsteady effects of shock wave induced separation

Coordinated by IMP PAN




Interaction types considered in UFAST:

Transonic interaction Nozzle flow Oblique shock reflection

Possibly wide representation of characteristic physical cases

of shock wave boundary layer interaction




No. Participant name Short name

1 Institute of Fluid Flow Machinery, Gdansk PL

IMP PAN

2 CNRS Lab. IUSTI, UMR 6595, Marseille F IUSTI

3 ONERA: ( DAFE, DAAP) F ONERA

4 University of Cambridge, Dept. of Engineering UK UCAM-DENG

5 Queens University Belfast, School of Aero. Eng. UK QUB

6 Russian Academy of Science, Siberian Branch, Novosibirsk, Inst. of Theor. App. Mech. RUS

ITAM

7 Delft University of Technology, Aerodyn. Lab. NL TUD

8 INCAS, Romanian Institute for Aeronautics RO INCAS

9 University of Southampton, (SES) UK SOTON

10 University of Rome "La Sapienza", I URMLS

11 University of Liverpool, Dept. of Aero. Engin. UK LIV

12 NUMECA, Belgium, SME B NUMECA

14 Institute Mécanique des Fluides de Toulouse F IMFT

16 FORTH/IACM, Found. for Res. and Techn.-Hellas GR FORTH

17 Ecole Centrale de Lyon F LMFA

19 EADS-M, Deutschland GmbH Military Aircraft D EADS-M

20 Institute of Aviation, Warsaw PL IoA

Large tunnels

Large tunnels

SME

New partnerUkraine

CFD

IND




Observer Name

1 Rolls-Royce Germany, RRD

2 Dassault Aviation

3 Alenia

4 Ansys Germany GmbH

Observer Group




PM=638

Transonic interaction Channel flow Shock reflection




ALTERNATIVE SOLUTION TO AJVG ALTERNATIVE SOLUTION TO AJVG

STREAMWISE VORTICITY CONTOURS

MEMS – sub-layer device ~ δ1

Rod Vortex Generator




Full Aero-thermal Combustor-Turbine interactiOn ResearchFACTOR

Zastosowanie AJVG w gorącym środowisku

2010 - 2013




Laminar – turbulent transition effects

on unsteady behaviour

of shock wave induced separation

New project topic:

UFAST TUFAST SWIFT?Przemysł: Dassault Aviation, Rolls-Royce Germany,

AVIO, AIRBUS




THE PEOPLE PROGRAMME – Marie-Curie

Industry-Academia Partnerships and PathwaysSTA-DY-WI-CO (LMS Belgium - IMP PAN Poland)

STAtic and DYnamic piezo-driven StreamWIse vortexgenerators for active flow Control

Initial Training NetworksIMESCON

Innovative MEthods of Separated Flow CONtrol in Aeronautics

piotr doerffer transonic flow control by streamwise vortices. research offer

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