1972_AGARD-AG-167_Modern Method of Testing Rotating Components of Turbomachines

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AGARDograph No. 167

Modern Methods of TestingRotating Components of Turbomachines

Edited by

M.Pianko

NORTH ATLANTIC TREATY ORGANIZATION

DISTRIBUTION AND AVAILABILITYON BACK COVER

AGARD-AG-167

NORTH ATLANTIC TREATY ORGANIZATION

ADVISORY GROUP FOR AEROSPACE RESEARCH AND DEVELOPMENT

ORGANISATION DU TRAITE DE L'ATLANTIQUE NORD

GROUPE CONSULTATIF POUR LA RECHERCHE ET LE DEVELOPPEMENT AEROSPATIAL

AGARDograph No. 167

MODERN METHODS OF TESTING ROTATING COMPONENTS OF TURBOMACHINES

Edited by

M.Pianko

Service Technique Aeronautique4, Avenue de la Porte d'Issy

75996 Paris ArmeesFrance

( Report of a Study organized by the AGARD Propulsion and Energetics Panel and comprising papers,^N^ discussions and conclusions, from the meeting held at the Ecole Nationale Superieure de

^X. 1'ASronautique et de 1'Espace, Toulouse, France, 18-21 September 1972.

THE MISSION OF AGARD

The mission of AGARD is to bring together the leading personalities of the NATO nations in the fields ofscience and technology relating to aerospace for the following purposes:

- Exchanging of scientific and technical information;

— Continuously stimulating advances in the aerospace sciences relevant to strengthening the common defenceposture;

- Improving the co-operation among member nations in aerospace research and development;

- Providing scientific and technical advice and assistance to the North Atlantic Military Committee in thefield of aerospace research and development;

- Rendering scientific and technical assistance, as requested, to other NATO bodies and to member nationsin connection with research and development problems in the aerospace field;

- Providing assistance to member nations for the purpose of increasing their scientific and technical potential;

- Recommending effective ways for the member nations to use their research and development capabilitiesfor the common benefit of the NATO community.

The highest authority within AGARD is the National Delegates Board consisting of officially appointed seniorrepresentatives from each member nation. The mission of AGARD is carried out through the Panels which arecomposed of experts appointed by the National Delegates, the Consultant and Exchange Program and the AerospaceApplications Studies Program. The results of AGARD work are reported to the member nations and the NATOAuthorities through the AGARD series of publications of which this is one.

Participation in AGARD activities is by invitation only and is normally limited to citizens of the NATO nations.

Part of the material in this publication has been reproduceddirectly from copy supplied by AGARD or the author.

Published May 1973

621-135:621-253:621.001.4

Printed by Technical Editing and Reproduction LtdHarford House, 7-9 Charlotte St, London. W1P 1HD

PREFACE

Tests on turbine-engine elements have always been part and parcel of engine manufacturers'ways and means, and actively contributed to the development of power plants; consequently,such tests have become more and more important. As a matter of fact, with the increasingcomplexity of turbine-engines, one is greatly tempted to isolate difficulties and improve under-standing by separating the various stages of the study involved. For example, at a given timeof the process, the testing and development of a turbine-engine is divided into investigations onits various components, namely: air intake, compressor, combustion chamber, turbine, nozzleand regulating system.

However, .simultaneously, the increasing complexity of turbine engines leads to anaugmentation of the interaction phenomena taking place between the different components,as well as of phenomena considered up till now as secondary and accessory. For this reason,the more and more extensive use of tests on partial elements paradoxically raises heretoforeunsuspected difficulties of interpretation, understanding and utilization.

Therefore, a tendency has been observed among the various users to endeavor to derivethe greatest possible cost-efficiency from the data acquired through partial tests. In each field,different, not to say conflicting, investigation and interpretation methods have been initiatedby industrial firms as well as research institutes.

The observation of this fact has encouraged the AGARD Propulsion and Energetics Panel toundertake a survey and comparison of the various methods applied. The Panel concern inthis field may be regarded as coinciding with that of many users who, through their work andeffective participation, have expressed interest in, not to say the need of, exchanging andclarifying views in this specific area.

Les essais sur elements des turbomachines ont toujours fait partie de I'arsenal desconstructeurs, et contribue activement a la phase de developpement d'un propulseur; de cefait Pimportance de ces essais n'a fait que croitre. En effet au fur et a mesure que lacomplexity des turbomachines augmente, la tentation est grande d'isoler les difficultds, etd'ameiiorer la comprehension, en proce"dant a une separation des etudes. C'est ainsi que1'essai et la mise au point d'une turbomachine est a un certain moment divise'e en Petudede ses differents composants a savoir, entree d'air, compresseur, chambre de combustion,turbine, tuyere et systeme de regulation.

Cependant la complexity grandissante des turbomachines entraine en meme temps uneaugmentation de Pimportance des phenomenes d'interaction des differents composantsentre eux ainsi que de phenomenes considers auparavant comme secondaires ou annexes.Pour cette raison, Pusage de plus en plus repandu d'essais sur elements partiels se heurte,paradoxalement, a des difficultes insoup?onnees jusqu'alors, d'interpretation, decomprehension et d'exploitation.

On a done vu les differents utilisateurs tenter de rentabiliser au mieux les resultatsacquis par les resultats essais partiels. Dans chaque domaine, des methodes d'6tude etd'interpretation differents, voire contradictoires, ont vu le jour, tant chez les industrielsque dans des organismes de recherche.

Cette constatation a incite le Groupe de Travail "Propulsion et Energ6tique" dePAGARD a entreprendre Petude et la confrontation des differentes methodes utilises.II est permis de penser que la preocupation du Groupe de Travail et aliee au devant decelles des nombreux utilisateurs, qui par leur travail et leur participation efficace ontmanifeste Pinteret sinon le besoin d'echanger et de clarifier les idees dans ce domaine.

L'Ingenieur en Chef PIANKO MarcPresident du Comite d'Organisation

iii

PANEL STUDY GROUP MEMBERS

M. Ping, en Chef Air M.Pianko (Study Leader)Service Technique AeronautiqueParisFrance

M. le Professeur J.ChauvinInstitut von Karman de Dynamique

des FluidesRhode-St-GeneseBelgium

J.F.ChevalierSNECMA VillarocheMoissy-CramayelFrance

Dr J.DunhamNGTEPyestock, FarnboroughHantsUK

Professor A.E.FuhsNaval Postgraduate SchoolMonterey, CaliforniaUSA

E.C.SimpsonAF Aero Propulsion LaboratoryWright-Patterson AFBOhioUSA

Professor Dr Ing.H.KiihlDFVLRInstitut fur LuftstrahlantriebePorz-WahnGermany

Dr Ing.H.StarkenDFVLRInstitut fiir LuftstrahlantriebePorz-WahnGermany

Dr P. van StaverenProjectgroep StromingsmachinesInstitute for Applied Research TNODelftNetherlands

TABLE OF CONTENTS

INTRODUCTION 1

VALUE AND USE OF CASCADE TESTS DATA 5

TESTING AND MEASURING EQUIPMENT FOR CASCADE TESTS 13TESTING TECHNIQUES FOR SUPERSONIC COMPRESSOR CASCADE

TESTS ON COMPRESSOR OR TURBINE STAGE(S) 19

TESTS ON A COMPLETE COMPRESSOR (OR TURBINE) 25REYNOLDS NUMBER EFFECTSCOLD TESTING OF TURBINES

COMPRESSOR STABILITY AND DISTORTION TESTS 31

GENERAL QUESTIONS 37

ANNEX - LIST OF COMPANIES AND INDIVIDUALS WHO PARTICIPATED IN THE STUDY 41

TABLE DES MATIERES

INTRODUCTION 3

EVALUATION ET ESTIMATION DES RESULTATS D'ESSAIS DE GRILLES 5

EQUIPEMENT DE MESURE ET D'ESSAIS POUR LES ESSAIS DE GRILLES 13TECHNIQUES D'ESSAIS POUR GRILLES DE COMPRESSEUR SUPERSONIQUE

ESSAIS SUR ETAGE(S) DE COMPRESSEUR OU TURBINE 19

ESSAIS SUR COMPRESSEUR (TURBINE) COMPLET(E) 25INFLUENCE DU NOMBRE DE REYNOLDSESSAIS DE TURBINE A FROID

ESSAIS D'HETEROGENEITE ET DE STABILITE DES COMPRESSEURS 31

QUESTIONS D'INTERET GENERAL 37

ANNEXE - LISTE DES SOCIETES ET DES PERSONNES AYANT PARTICIPE A L'ETUDE 41

INTRODUCTION

Under the aegis of the Propulsion and Energetics Panel (PEP) an ad hoc working group was set up in 1972 toconduct a study^on "Modern Methods of. Testing Rotating Components of Turbomachines". The membership ofthe Planning'Committee is:

- Professor J.CHAUVIN (Belgium)- J.F.CHEVALIER (France)- Dr J.DUNHAM (UK)- Prof.A.E.FUHS (USA)- Prof.H.KUHL (Germany)- I.C.A. M.PIANKO, Chairman (France)- E.C.SIMPSON (USA)- P.van STAVEREN (Netherlands)

-The approach of the Planning-Geirmrittee- was to draw up an inventory of the test methods for the study ofthe rotating components of a turbomachine, to describe them, and to bring out the essential ideas, the methods andconclusions applicable to the utilization and the interpretation of such tests.- \—/

_ —•- . " J»3s^~~~ - _.-- — -^-The study subjects^«=kr examined were:--

C..

lue and usefulness of cascade test data, ~^>-tfljesting and measuring equipment for cascade

Resting techniques for supersonic compressor cascade_s/ .Tests on compressor or turbine stage(s), - r • • - • - ' .Tests on a complete compressor (or turbine). „Compressor stability and distortion tests^ —J

Reynolds number effectstesting of turbines-; ^ — «===.-•=•*

^ — General questions ^* ,

The objective was to assess the advantages and usefulness of certain types of tests, the so-called "elementarytests" compared with the tests conducted on complete turbomachines. However, the scope of such a study seemedtoo wide to be covered in a single study. Therefore, the Planning Committee decided to limit the first stage to theaerodynamic aspect only (neither mechanics nor thermodynamics) of rotating components (no combustion, reheat,nozzle, or fuel control).

The Planning Committee considered that to fulfill such a task it. would be appropriate to bring together all thespecialists in the field to ensure an exchange of views and dialogue as broad as possible.

To reach this aim, the following procedure was used. First of all the Planning Committee drew up a question-naire which included for each study subject a number of detailed and specific questions. This questionnaire waswidely distributed and was directed to firms as well as research organizations, laboratories and universities. The listof all societies which answered the questionnaire appears in the Appendix. Each responder had also received acomplete set of the answers and therefore could make fruitful comparisons and analyses on their own account.

At the same time, the answers to the questionnaire were analyzed and investigated by experts appointed bythe Planning Committee. The list of people responsible for synthesizing the reports is given below:

- Value and use of cascade test data Messrs FABRI and PAULON

— Testing and measuring equipment for cascade tests+ Testing techniques for supersonic compressor cascade Dr H.STARKEN

- Tests on compressor or turbine stage(s) Messrs THIAVILLE and STOS

- Test on a complete compressor (or turbine)+ Reynolds number effects+ Cold testing of turbine Dr J.DUNHAM and Mr R.K.OLDHAM

- Compressor stability and distortion tests Professor G.OATES

- General questions I.C.A. M.PIANKO

The final stage of the study took place in September 1972 at Toulouse where discussions were organized amongthe specialists who had answered the questionnaire. During three consecutive days in an informal atmosphere, all the

people concerned could talk freely, comment, complete their answers and ask their colleagues present for anyexplanations. These exchanges of views, and comparisons were conducted with competence by the same expertsthat have prepared the synthesis reports on the questionnaire.

All participants appreciated the completeness and the variety of information given during this meeting wherethe specialists coming from various disciplines could very quickly find a common language which is something rare.To sum up, the different stages of the study were the following:

- a complete questionnaire

— varied and abundant answers

- a detailed investigation of all answers

— a free exchange of views amongst all specialists.

It seemed impractical to publish all documents issued during the study, and it would not be essential anywaybecause all those who took an interest in the study received all the papers relating to it and could follow itsprogress, phase by phase. Hence, the present publication includes only the general conclusions. Therefore thispublication consists of a resultant of the synthesis prepared on the basis of the answers to the questionnaires andincludes conclusions emanating from discussions in the final meeting.

During these discussions it became apparent that certain problems needed further investigation. In particular,it was decided to set up a Working Group chaired by Dr Hans Starken to initiate a detailed survey of cascademeasuring equipment. Therefore the work started on the "Modern Methods of Testing Rotating Components ofTurbomachines" will continue and other publications will follow the first conclusions presented in this document.

I N T R O D U C T I O N

Sous 1'egide du Croupe de Travail de I'AGABD "Propulsion et Energie" a e"t<S ore'e' en 1972 un GroupeAd Hoc charge1 de diriger une e'tude sur les "Me'thodes modernes d1 essais des parties tournantes des Turbomachines".Le Comit^ d1Organisation est compost de :

- Prof. J. CHAUVTN (Belgique)

- M. J.P. CHEVALIER (Prance)

- Dr J. DUNHAM (Grande-Bretagne)

- Prof. A.E. FUHS (Etats-Unis)

- Prof. H. KUHL (Allemagne)

- I.C.A. M. PIANKO, President (France)

- M. E.G. SIMPSON (Etats-Unis)

- M. P. Tan STAVEREN (Pays -Bas)

Le Comit^ d1Organisation s'est fixe1 comme progrannne de travail de faire 1'inventalre et la des-cription des moyens d1 essais mis en oeuvre pour I'e'tude des oomposants relatifs d'une turbomachine et de degagerles idees essentielles, les me'thodes et les conclusions conoernant 1'exploitation et I1interpretation de telsessais.

Les sujets d1etudes retenus sont I

- Evaluation et estimation des re'sultats d'essais de grilles,

- Equipement de mesure et d'essais pour les essais de grilles,

- Techniques d'essais pour grilles de compresseur supersonique,

- Essais sur e'tage(s) de compresseur ou turbine,

- Essais sur compresseur (turbine)complet(e),

- Essais d'he'teroge'ne'ite' et de stability des compreaseurs,

- Influence du nombre de Reynolds,

- Essais de turbine a froid,

- Questions d'interest general.

II s'agiseait de juger I'int rSt et I'utilite' de certains types d'essais dits "elementaires" encomparaison avec les essais effectues sur une turbomachine complete. Mais le domaine d'une telle e'tude a parutrop vaste pour 8tre explore" en une seule fois. C'est pourquoi le Comite" d1 Organisation n'a retenu dans une pre-miere phase que I1aspect a rodynamigue (pas de mdcanique nl de thermodynamique) des composants tournants (pas decombustion, rechauffe, tuyere, regulation).

Le Comit^ d1 Organisation a estime1 que pour mener a bien une telle e'tude il convenait d'dtablir uneconfrontation et un dialogue aussi large que possible entre tous les sp oialistes travaillant sur les problemes.Bour atteindre ce but, il a e'te' fait appel a la me'thodologie suivante. Tout d'abord le Comite" d1 Organisation abati un "questionnaire" qui, pour ohaque sujet d'etude, comprenait des questions nombreuses, ddtaille'es etprecises. Ce "questionnaire" a fait 1'objet d'une tres grande diffusion. II a ete fait appel aussi bien a desIndustriels, qu'a des Organismes de Reoherches, d1Etudes, des Laboratoires et des Uuiversit s. La liste de toutesles Soci t s ayant rdpondu au questionnaire figure a I1 Annexe. Chaque Soci td ayant repondu au questionnaire arecu toutes les r ponses et a pu ainsi, pour son propre compte, proceder a des comparaisons et analyses fruc-tueuses.

Parallelement, les r ponses au questionnaire ont e'te' analyse'es et e'tudie'es par des experts de'si-gn s par le Comit d'Organisation. Void la liste des personnes ayant effectue" oe travail de synthese :

- Evaluation et estimation des resultats d'essais de grilles MM. FABRIet PAULON

- Equipement de mesure et d'essais pour les essais de grilles+ Technique d'essais pour grilles de compresseur supersonique Dr HANS

STARKEN

- Essais sur etage(s) de compresseur a turbine MM. THIAVTLLEet STOS:

- Essais sur compresseur (turbine) complet(e)+ Influence du nombre de Reynolds+ Essais de turbines a froid ^ D™HAM

M. R.K. OLDHAM

- Essais d'heterogeneite et de stabilite" des compresseurs Prof. Gordon OATES

- Questions d'interet general I.C.A. PIANKO

Enfin la phase ultime de 1'etude eut lieu a Toulouse en Septembre 1972 oil des discussions furentorganisees entre les specialistes qui ont r^pondu au questionnaire. Pendant trois journ es consecutives, dansune ambiance d'ou tout formalisme etait exclu, tous les interess s purent s'exprimer librement, commenter, com-pieter leurs reponses et demander tous les eclaircissements et toutes explications a leurs collegues presents.Ces echanges de vue, ces confrontations furent dirige's avec competence par les mimes experts qui ont dejaeffectue les rapports de synthese sur le questionnaire. Tous les participants ont appre'cie' la richesse et lavariete des enseignements degages au cours de eette reunion ou fait pare, des specialistes venus des horizonstres divers ont pu rapidement trouver un langage oommun.

En resume, les diffdrentes phases de 1'etude ont ete t

- Un questionnaire complet,

- Des reponses variees et abondantes,

- Une exploitation detainee de toutes les reponses,

- Un libre echange de vue entre tous les specialistes.

II paralt exclu de publier tous les documents emis au cours de I1etude. Une telle publicationn'apporterait d'ailleurs rien d'essentiel car tous ceux qui se sont interesses a 1'etude ont pu, phase parphase, en suivre I'avancement et ont recu tous les documents s'y rapportant.

La presente publication ne comporte done que les conclusions generales. Elle est une resultants dessyntheses prepares a partir des reponses au questionnaire et des conclusions et idees qui se sont d4gagees aucours des discussions de la reunion finale.

Lors de ces dicussions il a ete constate que certains problemes meritaisnt un supplement d'etude.II a ete en particulier decide de creer un groupe de travail preside par Dr. HANS STARKEN charge d'etablir unrapport detailie sur les equipements de mesures dans les grilles. Le travail entrepris sur les "Methodesmodernes d'essais des parties tournantes des Turbomachines" continue done et d'autres publications suivrontles premieres conoluaions presentees ici.

EVALUATION ET ESTIMATION DES RESULTATS

D'ESSAIS EN GRILLES

Question n° 1 - Pensez-vous que les renseignements apportes par les essais de grilles peuvent s'appliquer auxecoulements tri-dimensionnels (point de calcul) ?

- directement

- aveo de petitea corrections (preoiser lesquelles)

- aveo des corrections importantes (preciser lesquelles)

- seulement en tant que guide qualitatif, pour comparer des profile par exemple.

Reponse i Les diverses reponses a cette question traitent aussi bien du subsonique que du supersonique etaussi bien des compresseurs que des turbines.

Les tendances qui s'en degagent sont les suivantes :

- en subsonique, 1'accord est quasi general pour une application directe aux turbines et auxcompresseurs (a la rigueur quelques petites corrections pour ces derniers).

- en supersonique, pour les turbines, 1'application reste possible avec des corrections assezimportantes. Par centre, les avis sont assez partages en ce qui conoerne les compresseurs.

La majorite pense qu'une transposition est possible avec des corrections plus ou moins impor-tantes sur les pertes et la deflexion selon que 1'on respecte AVDR - M - Re et le niveau deturbulence et si les effets radiaux sont faibles.

Pour environ 15 $ des reponses, les renseignements obtenus sont plutfit qualitatifs.

En ce qui concerne les corrections a appliquer, chacun a sa methode propre derive de sonexperience.

Commentaire : Aucune philosophic certaine n'a pu §tre degagee de la discussion. Certains experimentateurs onttrouve un recoupement correct entre essais en grille et sur compresseur, mais surtout dans ledomaine subsonique. En supersonique on ne dispose pas encore de suffisamment de donn es surmachine.

Question n" 2 - Pensez-vous que les renseignements apportes par les essais de grilles sont utiles pour les cas===== nnc)rg adaptation" ?

- dans 1 'affirmative, comment utilisez-vous les resultats d'essais ?Quelle est la plage de variation autour du point de oalcul ?

- dans la negative, pourquoi pensez-vous que oes renseignements sont inutiles ?

Reponse : Oui a une grande majorite, a condition qu'il n'y ait pas de decollement et que les effets ra-diaux soient faibles.

Toutefois, la plage de variation est assez limitee et les essais de grilles ne permettent pasde definir avec precision la limite de decrochage.

Commentaire i En general, on considere que les essais de grille ne peuvent donner que le point de fonctionne-ment et une petite plage autour de ce point et que pour le deorochage, les essais de grillesont sans valeur. Cependant, certains constructeurs utilisent les essais de grille pour 'pla limite du domaine de pompage.

eation n° 3 - Dans quelle mesure pouvez-vous appliquer a des composants tournants des valeurs numeriquesacquises sur des grilles ?

Reponse : oeneralement, la reponse est favorable avec certaines restrictions, deja clteea dans la r<5ponsen° 1.

De plus, le probleme de la definition de la coupe se pose et il y a, sur machine, une centrlfu-gation de la couche limite, des effets radiaux et des effets secondaires qui ne sont pas prisen oompte dans les essais de grilles. D'autre part, la charge llmite est plus eievee sur rotorqu'en grille.

Pas de commentaire

Question n° 4 - Comment faites-vous intervenir 1'influence de la variation du rapport des vitesses axiales~ sur les pertes, la rotation de 1'ecoulement et la charge limite ?

Repoase : Les methodes utilises sont tres variables.

- resultats d'essais a ATZDB variables,

- corrections empiriques,

- facteur de diffusion surtout pour les pertes,

- theories de Horlock - Acosta - Hawthorne - Stark - Schlichting - Jans Jansen,

- formule de Ohuonu.

Enfin la question est m8me posee de savoir si cela a un sens de restituer en grille et surmachine, le mime AVDR.

Commentaire t II ne faut pas sousestimer le role du rapport des vitesses axiales (AVR) ou le rapport duproduit vitesse axiale par densite (AVDR). II est tres important de restituer ce rapport dansles essais de soufflerie.

Question n" 5 - Utilisez-vous le concept du rapport des vitesses axiales ou celui du rapport masse volumique/vitesse axiale ?

Comment les definisaez -vous et !•• •esurez-vous ?

Reponse : De facon quasi generale, on utilise

AVR en subsoniqueet AVDR en supersonique

Vx etant la vitesse axiale.

La mesure est effectuee par Bondage sur un pas a mi-envergure.

Commentaire t On n'a pas pu trancher si le rapport des vitesses axiales est suffisant en fluide compressible,car il est caracteristique des effets secondaires, ou s'il faut utiliser le rapport du produitvitesse axiale par densite, caracteristique de la contraction de la veine.

Question n" 6 - Utilisez-vous les resultats des essais de grilles pour evaluer les effets d'un ecoulementsecondaire et, si oui, de quelle maniere ?

Reponse : Les avis sont tres partages.

- une partie n'utilise pas les resultats d'essais de grilles,

- une autre partie fait des mesures au voisinage des parois avec une couche limite amontd'epaisseur fixe ou reglable. Correlations de pertes,

- une troisieme partie pense que c'est interessant et plus correct pour des stators essayes engrille annulaire,

- enfin le problems est aussi etudie theoriquement.

Commentaire : II est tres souhaitable, de 1'avis general, que ce probleme des effets secondaires soit appro—fondi.

Question n° 7 - Ressentez-vous le besoin d'une recherche detaillee sur les ecoulements secondaires ? Si oui,==== comment faites-vous cette recherche ?

Reponse : Oui a 1'unanimite.

Le probleme est surtout interessant dans le domaine compressible et cette recherche est con—duite de fa?ons variees :

- mesures en grilles au voisinage des parois avec ou sans jeu, avec couche limite d'epaisseurvariable ou non,

— visualisations diverses,

- mesures au fil chaud,

- mesures sur machines tournantes avee influence du jeu,

- analogic hydraulique,

- etude theorique menee en liaison avec les essais dans certains cas.

Commentaire : Voir question n° 6.

Question n° 8 - Quelles mesures prenez-vous ou considerez-vous necessaires en ce qui concerne la rugosite etles decollements locaux de couche limite, ou le controls de la turbulence ou des couches limi-tes de paroi ?

Les couches limites de paroi sont aspires si cela se revele necessaire. Quant a la turbulence,elle est simuiee autant que possible, mais certains pensent que c'est un probleme tres diffici-le.

Commentaire : II semble que le dedenchement artificiel de la transition risque de modifier 1'ecoulement defacon differente et conduit parfois a des resultats errones.

Question n° 9 - Comment utiliser les resultats d'essais de grille pour calculer le domaine de fonctionnementd'une turbomachine ?

Reponse : Reponaes generalement negatives, sauf pour quelques-unes qui le jugent possible au voisinagedu point d'adaptation et de preference pour les turbines. Quelques methodes de calcul existentou sont en etude. Certains constructeurs precisent la limite du pompage grftce aux essais degrille.

Commentaire : Pas d'avis compiementaires.

Question n° 10 - Comment definissez-vous les conditions "moyennes" d'entree et de sortie de I'ecoulement, telleeque les deflections, la pression, la vitesse, etc. (ponderation en section, ponderation endebit-masse ...?)

Reponse : Par ponderation en debit a la grande majorite.

Commentaire » Le probleme de la presentation des conditions moyennes a ete discute en detail. Les resultatsles plus consistants sont obtenus par calcul en aval de la grille des conditions fictives queI'on obtiendrait par uniformisation du melange a section constante. II faut pour cela teniroompte des equations de conservation de 1'energie, de la quantite de mouvement et de debit.

II a ete recommande que desormais dans toute publication figurent les resultats apresmelange complet a section constante.

Question n" 11 - Comment d finissez-vous la section d'aube equivalents a la section de grille sur la machine ?~ (sur une surface eylindrique ou conique, ou sur une ligne de courant).

Reponse i Sur une surface eylindrique en subsonique et sur une surface conique ou mieux sur une ligne decourant pour les machines avanc es.

Commentaire i II semble bien que la definition de la coupe equivalente & une grille d'aube d pende du labora-toire. La plupart du temps on se contente de definitions simples telles que coupe 'conique.

Question n° 12 - Pennez-vous qu'il convienne de simuler convenablement les influences suivantes t

- allongement,

- jeu,

- epaisseur de la couche limite a 1'entree,

- derapage de la couche limite a 1*entree,

10

Dans I1affirmative, comment les simulez-vous ?

Reponae : Oui pour 1"ensemble en general; mais si 1'allongement peut 8tre assez faoilement simuie, iln'en est pas de m&me pour les autres points et surtout, il n'est pas certain que les resultatsobtenus soient transposables.

Commentaire i II faut rajouter un faeteur compiementaire i le gradient axial de pression.

Question n" 13 - Pensez-vous que les essais de grille puissent apporter des renseignements utiles en ce qui" concerae les ecoulements instationnaires ? (coefficients d'amortissement aerodynamique).

Reponse : Gene"ralement oui. Mais la mise en oeuvre doit Stre correcte.

Commentaire : II existe mSme des souffleries ou des ecoulements instationnaires contrOies ont ete realises.

Question n" 14 - Utilisez-vous les essais de grille pour evaluer 1"influence aerodynamique du refroidissement~ des aubes de turbine ?

R^pcnse : Oui en general.

Pas de commentaire.

Question n° 15 - Quelles sont les meilleures techniques de visualisation d'un ecoulement ?Quels sont les renseignements que nous apporte la visualisation ?

Reponse : Les techniques sont tres variees selon les buts recherches et les conditions d'utilisation.

On peut diatingusr pratiquement deux groupes :- pour les faibles vitesses t la furnee, les fils de laine ou de soie, la peinture ou les divers

melanges, 1*analogic hydraulique avec poudres ou bulles, etc...- pour le supersonique i la methode des ombres, la strioscopie, 1'holographie, 1'interferome-

trie, la peinture, I'analogie hydraulique, etc...

La visualisation est un atout important dans la comprehension des phenomenes puisqu'elle permetde localiser des de'collements, des tourbillons, des zones de transition, des ondes de chocs,etc...

Commentaire : La technique de 1'holographie laser est tres attrayante, elle n'est pas encore au point. II estrecommande de pousser le developpement de cette technique.

Question n° 16 - Etes-vous satisfait des correlations experimentales de grilles disponibles (N.A.S.A., Ainley-======= Mathieson, eto...) ?

Reponse : Oui pour les grilles correspondantes et dans le domaine ou elles ont ete etudie'es.

Rarement pour d'autres grilles et surtout aux grandes vitesses.

Un gros travail reste a fairs dans oe domaine.

Commentaire : La standardisation des formules de correlation est loin d'Stre obtenue. Utilis^es pour lesmdmes essais elles donnent g^n^ralement les mgmes resultats.

AGARDograph No.167Advisory Group for Aerospace Research andDevelopment, NATOMODERN METHODS OF TESTING ROTATINGCOMPONENTS OF TURBOMACHINESEdited by M.PiankoPublished May 197352 pages

Under the aegis of the Propulsion and EnergeticsPanel (PEP) of AGARD, a study was conducted ontest methods and means for turbomachine rotatingcomponents. More than forty companies fromBelgium, France, Germany, the Netherlands, theUnited Kingdom, the USA; engine manufacturers

P.T.O.

AGARD-AG-167621-135621-253621.001.4

TurbomachineryRotorsRotor blades

(turbomachinery)Test facilitiesTest equipment

AGARDograph No. 167Advisory Group for Aerospace Research andDevelopment, NATOMODERN METHODS OF TESTING ROTATINGCOMPONENTS OF TURBOMACHINESEdited by M.PiankoPublished May 197352 pages


P.T.O.

AGARD-AG-167621-135621-253621.001.4





P.T.O.

AGARD-AG-167621-135621-253621.001.4





P.T.O.

AGARD-AG-167621-135621-253621.001.4



as well as institutes or research organizations could take advantage of this oppor-tunity to compare their experiences, and their methods and make a fruitfulexchange of views. The present publication gives the synthesis and the generalconclusions of the study.

This AGARDograph comprises papers, discussions and conclusions from themeeting held at the Ecole Nationale Superieure de 1'Aeronautique et de 1'Espace,Toulouse, France, 18-21 September 1972.


This AGARDograph comprises papers, discussions and conclusions from themeeting held at the Ecole Nationale Superieure de I'Aeronautique et de 1'Espace,Toulouse, France, 18-21 September 1972.


This AGARDograph comprises papers, discussions and conclusions from themeeting held at the Ecole Nationale Superieure de I'Aeronautique et de 1'Espace,Toulouse, France, 18—21 September 1972.



11

Question n" 17 - A la lumiere des preoccupations relevees dans les questions precedentes, quelle est votre opi-nion concernant les avantages ou la necessite des essais en :

- grilles planes,

- grilles annulaires fixes,

- grilles annulaires mobiles ?

Grilles planes : Elles sont generalement jugees utiles, parce que commodes, peu chores, et qu'elles permettent 1'etude de phenomenes de base et de parametres particuliers. Ceei, malgri lesproblemes de transposition aux machines des resultats obtenus.

Grilles annulaires s Bien que moins experimentees, ees grilles suscitent neanmoins de 1'interStpour 1'etude de stators et des ecoulements secondaires principalement. La gSometrie se rappro-che de la machine et la periodicite est meilleure qu'en grille plane.

Grilles annulaires mobiles i Elles sont encore moins courantes et lea avis sont partages,

- jugees cheres ettrop proches de la machine pour ne pas preferer directement celle-ci,

- interessantes car eompiementaires des autres montages : eentrifugation de la couche limite,possibilite d'etudier les effets de jeu. De plus, grandes facilites pour couvrir un largedomaine de fonctionnement.

Sur I1ensemble du probleme grilles, on peut toutefois remarquer que de nombreux laboratoiressouhaiteraient pouvoir disposer des trois types d1installation.

Commentaire : Le choix du type de montage depend de criteres varies.Unanimement les grilles planes sont considerees comme les plus simples, mais des reserves doi-vent Stre faites sur la maniere dont les resultats peuvent representer 1"ecoulement dans lesturbomachines.

13

TESTING AND MEASURING EQUIPMENT FOR CASCADE TESTS

TESTING TECHNIQUES FOR SUPERSONIC COMPRESSOR CASCADE

15

TESTING AND MEASURING EQUIPMENT FOR CASCADE TESTS

The investigation, underlying this paper, has shown that cascade results are still used in the design ofaxial flow turbomachines. Therefore, measurements in cascade wind tunnels are necessary and consequentlythey are performed in various research institutes as well as industrial companies all over the world.

The main difficulties associated with these measurements can be separated under two head lines.

Firstly the problems connected with the correct simulation of the desired flow pattern, which -is mainlya function of the special wind tunnel design. Secondly the problems which arise with the measurement ofthe flow properties themselves, as for instance the static pressure, flow angle etc.

The discussions in the following chapters do not cover all problems related with the experimental cascadework. They have been selected, because they are important and different opinions do exist.

Cascade_Flow_Simulation

Theoretically the two-dimensional cascade requires an infinitely large number of blades. This cannot berealized in an actual wind tunnel. Generally the size of the test section is prescribed rather by theavailable air supply system than by aerodynamic requirements and consequently the blade number is fixedby the given cascade geometry as stagger angle and pitch-chord ratio. A slight variation in blade numberis possible by using different blade chord length, but this may lead to other difficulties like bladebending or problems with the measuring equipment, which has to be reduced in size.

A general investigation of this fluid mechanical problem has not been performed so far. Only in the spe-cial case of supersonic velocity at inlet or exit of a cascade some theoretical work has been published.It shows that in the supersonic flow field only three blades are required to get a periodic flow condi-tion. But this result may be quite different at subsonic or transonic velocities, where the boundaries doinfluence the whole flow field.

Because of this lack of basic knowledge of the cascade flow at a limited number of blades, the opinionsof the different specialists on this field do vary considerably. The quoted minimum blade number rangesbetween 5 and 15 blades for subsonic cascades and 3 to 9 blades for supersonic cascades. S.N.E.C.M.A.even proposes two blades to be sufficient if only investigations of the boundary layer behaviour and ofthe pressure distribution are of interest, whereas more than 1C should be used if losses are to be measu-red. In general the minimum number of blades, required for good periodicity, can be reduced by applica-tion of an extensive control mechanism, using flaps with suction or blowing, tailboards etc. and is,therefore, also a function of the specific wind tunnel design itself.

Especially for supersonic cascade wind tunnels the test section size is normally prescribed and the ques-tion is not, how many blades are necessary, but, how is it possible to improve or to get a reasonableflow periodicity with a given number of blades.

As already mentioned, in the case of supersonic inlet flow Mach number some theoretical results areavailable today and it is therefore possible to improve the periodicity by application of this know-ledge.

In the calculation of the supersonic inlet flow field of a cascade, the condition of periodicity issatisfied by the assumption of a uniform flow field infinitely far upstream. Therefore, in a two-dimensional cascade wind tunnel a supersonic flow field with constant properties has to be establishedupstream of the blades by using for instance a Laval-nozzle. The resulting periodicity behind the firstleading blade is perfect, or of the quality of the nozzle flow, when the axial component of the inletMach number is supersonic. At subsonic axial condition, the periodicity is influenced by the limitednumber of shock waves upstream of the cascade. But for normal blade shapes this influence is less thanthe measurement accuracy. The possible improvement of periodicity by a large number of blades is verysmall.

According to the "unique incidence1' relation, the inlet flow field adjusts itself at a certain Mach num-ber and flow angle, which is only a function of the nozzle Mach number, cascade geometry, and cascadeposition. In order to simplify the measurements, it is also possible to set the cascade at an inlet flowangle, determined by an analytical calculation. Under this condition, inlet Mach number and nozzle Machnumber are equal.

At transonic velocities (0.9 < M] < 1.2), slotted or porous walls with suction systems have to be addedto the cascade wind tunnel in order to avoid or to reduce shock wave reflections and choking of the windtunnel.

The above ideas are strictly correct only under the assumption of frictionless flow. In real flow theinlet conditions may be influenced by the side wall boundary layers. This objection has been raised bythe O.N.E.R.A. because some measurements do indicate such an influence, but it has not been proved sofar.

Cascade measurements at the DFVLR show, that the side wall static pressure, upstream of the cascade is infact influenced by the boundary layers and can be wrong, especially at high inlet Mach numbers with heavypressure gradients. But the measurements of the blade surface pressure distributions at mid span do notshow such an influence in the entry region of the blade passages. Whether these different results dependon the cascade geometry, including the aspect ratio, or not has to be investigated in further tests, also,

16

if it is possible to reduce any boundary layer influence by applying side wall suction, upstream of thecascade. But in this case an increase in secondary flow within the cascade due to the laminar side wallboundary layers may be suspected.

The same questions about the periodicity improvement of the inlet flow field of a cascade do exist in theexit flow field too. If the downstream Mach number is supersonic, it is again possible to solve this pro-blem at least qualitatively by using the inviscid flow model. Again the result is, that a uniform super-sonic flow field has to be provided, now in the exit flow downstream of the first blade, in order toestablish a limited area of periodicity. The accuracy of this method depends on the magnitude of theviscous effects in the wake region of the blades, and must therefore be determined experimentally. Butno results are available so far on this subject.

Even more difficult is the problem to obtain good- periodicity at the exit of a supersonic cascade, whenthe outlet Mach number is subsonic. Several methods have been tried in the past and it seems that thesolution depends very much on the special design of the cascade wind tunnel.

Basically two different constructions are possible at the exit of the cascade, either a dump diffusion ora guided exhaust. In the first case, a reasonable periodicity can be achieved by removing the air in theouter passages using a special exhaust system, and throttling the main flow far downstream of the dumpdiffusor by a valve. In the latter case, it is necessary to use tailboards, i.e. rotatable plates whichare fixed to the outer blades, in order to separate the main flow from disturbances in the outer passages.Thus it is possible to vary the back pressure without influencing the upstream cascade flow through theouter passages.

Unfortunately the throttle position has to be near the blade trailing edges, because otherwise the backpressure would be constant normal to the tailboards and not constant parallel to the circumferentialdirection of the cascade, as required for periodicity. On the other side, the downstream traversing planehas to have a similar position too and this may lead to some influence on the probe measurements. There-fore, a compromise has to be made between periodicity requirements and interference free flow.

All these difficulties, which arise because of the limited number of blades in a straight cascade windtunnel are avoided in the annular cascade, but at the expense of high cost and additional three-dimen-sional effects.

On the other hand the tendency in research work on straight cascades shows that more and more three dimen-sional effects are also introduced and measured in these facilities. Especially the axial velocity ratioor axial velocity density ratio is a very important parameter in axial compressor cascades. It has to betaken into account in all applications and comparisons of different measurements. For this reason, a pos-sibility to vary this parameter in a wind tunnel is useful, if not absolutely necessary.

Several methods have been proposed to attain this, like wall contouring and suction or blowing throughporous or slotted side walls. So far mainly the suction system has been applied, and there seems to be noinfluence of the porousity on the test results even with single slots. Some experiments, performed withcontoured side walls indicate that the measured flow conditions at mid span may be quite different fromthe design values. It is therefore not surprising that up to now all theoretical methods fail in thecorrect prediction of the axial velocity density ratio influence on cascade performance. Some theoriespredict for example, an increase in deviation with increasing axial velocity ratio, but the experimentalresults show the opposite tendency. Above this, no experimental data at all have yet been published aboutthe axial velocity influence on supersonic cascades.

In connection with the flow simulation, also the operating range of a cascade is of interest, because the-se data are useful to predict the off-design performance of the turbomachine. Therefore, in the subsoniccase the incidence is varied, if possible, between positive and negative stall and choking, whereby theinlet flow velocity is kept constant. However, in supersonic cascades with axial subsonic inlet velocity,the incidence is uniquely fixed by inlet flow Mach number and cascade geometry (unique incidence relation).Only at low supersonic velocities a slight variation of the inlet flow angle is possible. But generallyin supersonic cascade tests this flow angle is not an independent variable. It is replaced by the backpressure. In these cases, the operating range is defined by the range of back pressure, which can beachieved in the cascade. The limits are given by the open throttle condition, respective the minimumattainable back pressure, and by the spill point condition at maximum back pressure.

Thereby the cascade spill point is mostly defined as that point at which the inlet aerodynamic conditionsbegin to change with back pressure. This influence of the upstream flow by the back pressure occurswhen the normal shock waves in the blade passages reach the blade leading edges. Usually the spill pointcondition is detected by observation of the upstream side wall static pressure distribution and bySchlieren inspection. However, some results, especially at higher inlet Mach number, show that the sidewall static pressure indicates the spill point at a somewhat lower back pressure level than the Schlieren-picture. This is probable due to an influence of the side wall boundary layer, which allows an upstreamtravelling of the disturbances. These discrepancies can be avoided by using the blade suction surface pres-sure distribution for the determination of the spill point condition.

Cascade Flow Measurement

Basically there is no difference in the measurement of flow properties in a cascade wind tunnel from thatin a normal wind tunnel. But because of the strong gradients in flow velocity and flow angle, which dooccur in cascades, special designed probes are very often required. In consequence of the different expe-rience in various companies and research institutes also very different types of probes are used.

17

Besides single probes like cobra probes,Conrad probes, claw probes, wedge probes, Gottingen probes, NACA prism-type probes, AGARD static probes,Kiel probes, banjo probes, hot film and hot wire probes, also individual probes and rakes are applied.

In order to find out if there is any influence of the different probe types on the experimental testresults, a special committee has been proposed. This committee should not only investigate differentprobe-types and develop, if possible, standard probes for turbomachinery measurements but also specifythe related test procedures and data reduction methods. Past discussions on these subjects have shownthat a broad disagreement exists about the accuracy of flow measurements. Mainly the correct determina-tion of static pressures, flow angles, and in certain cases also total pressures (fluctuating flow, measu-rement at supersonic velocities) are difficult.

The measurement of static pressures in cascade wind tunnels can be performed in two ways. Either by staticprobes or by wall static taps. Under the assumption of complete two-dimensional flow, the static pressureat mid span and at the side walls should Be equal. In reality, however, only an approximate two-dimensio-nal flow condition can be achieved and the static pressure at mid span may Be different from that at theside walls. Consequently both measurements are very often compared. The agreement is good, or at leastsatisfactory, in most subsonic cascade measurements whereas in supersonic cascades larger differences dooccur. They are extremely high if the pressure gradients in the flow field are large as for instance inexpansion fans and shock waves. This is probably due to the existence of boundary layers, which can sepa-rate from the side walls and by this way enable a pressure balance in flow direction.

On the other hand, static probes have a limited accuracy too, and especially at transonic velocities thismay also result in large discrepancies. The quoted accuracies in pressure measurements range between0.02 % and 1 %, whereby the lower value should be attainable only at lower flow velocity. Similar restric-tions do apply to the flow angle measurement, where the accuracy is specified between 0.1° and 1°. Thevalue of 0.1° has been subject to additional questions, whether this includes all possible errors asintroduced by calibration, zero setting, mechanical backlash etc.. or not.A complete agreement could notbe obtained between the different specialists. However, some understanding was possible that at higherflow Mach numbers an accuracy of 0.5° to 1° should be much more realistic.

Again, different from flow investigations around single bodies, where force measurements are generallyperformed by balance systems, in cascade measurements lift and drag coefficients are determined indirectlyby integration of the flow quantities. This means that in two reference planes upstream and downstream ofthe cascade the flow properties have to be measured as accurate as possible.

Although the actual position of these planes can theoretically be chosen at will, there are several limi-tations because of the limited length of the cascade in a wind tunnel and because of the influence of thesurrounding wall boundaries.

The upstream flow is generated by a nozzle and it is normally assumed that the flow direction is parallelto the nozzle axis. However, the experience in subsonic cascade research showed that this is not alwaystrue and differences of several degrees in inlet flow direction upstream of the cascade may occur. There-fore an inlet flow traverse is desirable in subsonic cascades, although it is not always performed due tothe additional effort. But at least the inlet static pressure distribution has to be checked at the sidewalls. The traversing plane is generally taken a quarter to one chord length upstream of the blades where-by the distance of one chord is mostly preferred.

At supersonic inlet velocities a cgrresponding flow survey is extremely difficult because of flow angleand velocity gradients. In addition these measurements have to be performed in a triangular region behindthe leading edge shock or expansion wave of the first blade.

With increasing Mach number this region reduces in size until it vanishes completely at axial sonic inletflow conditions. However, as already mentioned before, there exists an unique relationship between theupstream supersonic nozzle flow and the inlet flow of the cascade. If this relationship is known, it ispossible to derive the cascade inlet conditions from the measured nozzle flow. On the other hand thismethod can only be applied at purely supersonic and not at transonic velocities.

Very similar considerations do apply to the reference plane in the exit flow field, but with the additio-nal difficulty of strong velocity gradients already existing at subsonic velocities. In itself it wouldbe advantageous to have the traversing plane at the blade trailing edges, because the wall influenceswould be minimized. Unfortunately there is no probe available today that gives correct readings of staticpressure and flow angles in a flow field with heavy gradients. It is therefore necessary to move the probeposition downstream and generally one quarter to one chord length behind the blades is assumed to givegood results. Again the distance of one chord length is preferred.

Besides the probe position also the data reduction method is important when different test results have tobe compared. This applies especially to the measurements of blade profile losses, which are generally pre-sented as averaged loss coefficients. These mean values are theoretically independent from the probe posi-tion if the equations of state (continuity, energy, and momentum) are used, and therefore all measurementsshould be published in this. kind. Although it has not been possible so far to verify completely the inde-pendency of the results from the probe position these date should be preferred Because they include alsothe mixing losses.

19

ESSAIS SUR ETAGE(S) DE COMPRESSEUR OU TURBINE

21

ESSAIS SUR ETAGE(S) DE COMPRESSEUR OU TURBINE

INTRODUCTION

L'etude rSalisee en congres s'appuyait sur les reponses Scrites faites par les participants a un ques-tionnaire. La synthese de ces reponses, point de depart de la discussion, faisait 1'objet d'un documentS.N.E.C.M.A. n° 3899/72 du 25.8.72 et ne sera pas reproduite ici. Les sujets traites en reunion plSnierepar les specialistes se rattachent en fait a deux themes principaux :

a) Adaptation et empilage des etages

b) Methodes et moyens de mesure

2. GENERALITES

Les etudes de plus en plus poussees qui precedent et accompagnent le dSveloppement des turbomachinesmodernes, en particulier dans le domaine des moteurs d'aviation, les recherches fondamentales et appli-quees sur lesquelles elles s'appuient font appel a une large gamme d'essais qui va du montage precis etpartiel en laboratoire jusqu'a 1'experimentation d'ensembles complets a 1'echelle industrielle. Apresles essais en grilles qui font 1'objet des deux premieres sessions, la session 3 aborde les partiestournantes proprement dites : compresseur ou turbine. Ces parties tournantes comportent souvent plu-sieurs etages et leurs essais mettent en oeuvre une instrumentation complexe et des puissances d'ins-tallation considerables. D'autre part, les essais effectues ne consistent pas seulement a mesurer eta amSliorer les performances globales au point d'adaptation mais aussi a rechercher une plage de fonc-tionnement de plus en plus large et a sender les caracteristiques internes de 1'ecoulement pour unemeilleure comprehension des phenomenes physiques (aerodynamique, thermodynamique, thermique) et de leurinteraction avec la technologic, ainsi qu'une amelioration des methodes de calcul.

3. ADAPTATION ET EMPILAGE DES ETAGES

L'adaptation relative des etages fait appel a la connaissance des caracteristiques propres de chaqueStage dans 1'ensemble tournant et les criteres de cout et de facilite de mise en oeuvre poussent 1'ingS-nieur de developpement a rechercher le montage le plus simple. L'essai d'un Stage isole facilite d'ail-leurs 1'introduction de systemes de mesures complexes et fragiles et les verifications experimentalesde schemas ou de methodes theoriques imposent aussi au chercheur la methode de 1'essai partiel. Aussin'est-il pas Stonnant de voir les specialistes se poser la question de la reprSsentativitS des essaisd'un etage ou meme d'une roue isolee.

Quelles que soient les reponses apportees, une premiere conclusion est aussitot a tirer des debats :il n'y a pratiquement personne pour nier 1'interet et mime la necessitS des essais partiels (etagesisolSs).

En fait, plusieurs methodes d'approche du probleme de leur representativite sont apparues dans la dis-cussion sans que 1'assemblSe puisse vraiment trancher sur les merites respectifs de chacune.

3.1. Absence ou faiblesse de 1'interaction entre etages

Un premier cas relativement simple est evidemment celui ou les conditions aux limites de 1'Stage isolS(ecoulement amont et Scoulement aval) ne peuvent etre alterees lors de son introduction dans 1'ensemblemulti-Stages. Du point de vue rigoureusement scientifique, cette situation n'est en fait pas possible :Stat des couches limites amont, turbulence, sillages fixes ou mobiles, Schanges avec 1'extSrieur sontautant de conditions qui ne peuvent etre remplies simultanement pour assurer la yalidi'tS totale de1'essai partiel. Mais, surtout en ce qui concerne les performances globales Ctaux de. compression ou dedStente, dSbit, rendement) trois exemples ont pu etre donnes d'une bonne reprSsentativite de 1'essaipartiel du fait de faibles interactions.

3.1.1. Compresseurs centrifuges

Le passage d'un rouet a 1'autre par 1'intermediaire d'un diffuseur (parfois supersonique), d'une voluted'echappement a tres faibles vitesses suivie d'une volute de captation accelerante -ces deux dernidresStant quelquefois reliSes par un systeme de brides et tuyaux-, limite les phSnomenes d'interaction.L'exemple de la realisation d'un compresseur a 3 etages dont on a pu prevoir le fonctionnement, memehors adaptation, a1 partir des essais d'un Stage isolS est venu appuyer cette these. ^~^

Un certain nombre de machines industrielles du meme type (pompes, compresseurs centrifuges subsoniquesa taux de compression relativement faible) semblent pouvoir relever de cette mSthode experimental.

3.1.2. Premier etage d'une turbine ou d'un compresseur axial

Sur ce cas particulier, 1'unanimite des participants n'a pu se faire que pour les performances au pointd'adaptation. En ce point les specialistes sont d'accord pour affirmer la faible influence vers 1'amont

22

d'un Stage correctement adapts ajoutS a 1'arriere. II est cependant necessaire de representer au mieux,dans 1'essai partiel, I'Scoulement le plus probable 5 1'entree. Toutes les conditions ne sont pas facile-ment reproductibles. Par exemple 1'effet de 1'Stat des couches limites amont sur les pertes secondairesdans 1'Stage fait encore 1'objet de recherches alors qu'il n'est pas toujours exactement semblable 3 laconfiguration moteur. Les essais avec hetSrogenSitS de pression amont sont realises. Us sont encoreplus nScessaire si 1'essai de 1'Stage isolS cherche les performances hors adaptation, specialement lalimite de pompage du compresseur qui peut etre modifiee par adjonction d'etages aval (stabilisation),ou par hStSrogeneitS amont.

3.1.3. Caracterisation d'un fan sans Stages I.P.

Dans ce cas, voisin du prScSdent, les specialistes pensent egalement que les interactions peuvent etrefaibles sous certaines conditions : Stude limitee au flux secondaire, vannages sSpares pour reglage dutaux de dilution, separation des flux suffisamment SloignSe du FAN. L'interpretation des essais estdelicate en ce qui concerne le domaine de fonctionnement.

3.2. Etage intermediate ou arriere d'une machine axiale

II s'agit du cas le plus general ou la reprSsentativitS rigoureuse ne semble pas possible. Effets vis-queux et instationnaires, deplacements radiaux du dSbit avec courbure des lignes de courant, tous cesphSnomenes baptisSs gSneralement "interaction" obligent 1'experimentateur d'un Stage isolS a tenter dereprSsenter, en amont et en aval, I'Scoulement le plus probable attendu dans la machine complete.

Certains organismes signalent a ce propos leurs Studes de 1'influence des sillages amont, de la turbu-lence, de la couche limite et des ecoulements secondaires (jeu radial par exemple) sur les performancesde grilles planes ou annulaires, d'etages. La reproduction artificielle d'un certain taux de turbulencesans distorsion du champ de pression stationnaire est possible, mais dSlicate.

Le probleme le plus difficile parait etre la prSvision du champ complet de performances d'une machinemulti-Stages. Certains domaines peuvent etre simplement prSvus par 1'Stage isole : en survitesse, bio-cage du premier Stage, pompage du dernier.

Mais la prSvision complete fait apparaitre nettement deux Scoles, deux mSthodes de travail.

3.2.1. Methode de 1'empilage

Certains specialistes, principalement sur machines de faible puissance, basent toute leur prSvision sur1'essai fouillS d'un Stage isolS. Cet essai est cependant suivi par celui de groupes d'Stages. La prSvi-sion s'effectue ensuite par "empilage" des performances partielles, elles-memes corrigSes empiriquementpour tenir compte des interactions.

Les coefficients empiriques de correction depassent la simple prScision des mesures et font appel a unegrande masse de rSsultats d'essais. Les corrections sont d'ailleurs sujettes £l variation selon le type demachine envisagS.

3.2.2. MSthode de 1'essai global

D'autres specialistes (principalement sur machines de grande puissance) indiquent avoir tente dans lepassS la mSthode prScSdente sans succes. Elle les a conduits finalement a effectuer 1'essai de la machinecomplete pour la mise au point et le dSveloppement. Leur mSthode consiste alors a relever directement, 31'intSrieur de 1'ensemble complet, les caracteristiques de chaque Stage (Cf.difficultSs de mesures auchapitre 4) et a chercher a les modifier jusqu'a obtention des performances globales desirSes.

4. METHODES ET MOYENS DE MESURE

La deuxieme partie de la discussion a ports sur le dSpouillement des essais et les moyens de mesure uti-lisSs.

4.1. Determination du rendement

Pour la dStermination du rendement,les mesures, soit des temperatures radiales et azimutales, soit ducouple, sont utilisSes suivant les installations. Pour permettre un recoupement possible des rSsultats,les,deux mesures sont souvent relevSes conjointement.

La mesure de la puissance sur 1'arbre est considSrSe en general comme plus precise : - 1 % sur le rende-ment pour 1 a 2 % avec les temperatures. Cette prScision varie Svidemment en fonction du niveau de varia-tion de temperature a mesurer et dSpend done du taux de compression ou de detente de la machine essayee.En effet, il est plus facile de connaitre les corrections a apporter au couple qu'aux temperatures ou a1'erreur de mesure proprement dite s'ajoute.dans certains cas.les effets de distorsion de I'Scoulementamont.

Par centre, les difficultSs de mise en oeuvre d'un frein ou d'un torsiometre font souvent preferer lamesure directe des temperatures. Un autre avantage de ce relevS est de permettre de connaitre en tout

23

point de fonctionnement de la machine la repartition radiale de la variation d'enthalpie.

Un grand nombre de systemes de mesure du couple sont utilisSs. Us ont leurs avantages et inconvSnientspropres selon le montage, la taille de la machine et 1'application envisagSe. Citons par exemple : tor-siometres a systemes optiques, Slectriques, magnetiques, jauges extensiometriques, couple-metres a jauges,a dynamometre, freins a eau, etc...

Quant a la determination de la pression totale a la sortie d'un Stage deux pratiques sont utilisSes :soit la mesure de 1'angle et des pressions statiques de parois, la pression totale se deduisant par conti-nuitS du dSbit, soit la mesure directe de la pression totale par exploration d'une sonde unique ou parpeignes de sondes multiples. Toutefois une orientation semble se dessiner en faveur des pressions totalesmesurees. En effet, beaucoup de constructeurs mettent en cause la valeur des pressions statiques deparois, en particulier il n'est pas possible dans de nombreuses machines de les placer dans des rSgionsnon perturbSes, comme dans les soufflantes a rapport de dilution Sieve ou dans les turbines entre lasortie de 1'aubage mobile et les bras de raccordement.

4.2. Moyens de mesure entre aubages

Dans le but de mieux connaitre I'Scoulement dans la machine, des explorations entre les aubages sontnScessaires. Toutefois tous les participants se montrent tres prudents quant a 1'interpretation desrSsultats. En effet, il est difficile d'Sviter toute interaction de la sonde avec I'Scoulement et enparticulier la prSsence de la sonde risque de modifier le blocage dans un aubage fixe et de crSer unStranglement dans 1'aubage mobile.

En general, les sondes sont telecommandSes, mobiles en profondeur et en orientation. Les explorationsfaites en plusieurs positions periphSriques permettent de couvrir un pas complet de 1'aubage fixe aumoins. Une tres grande variStS de sondes sont utilisSes, chacun mettant au point en gSnSral sa propresonde : sondes "a coin" associSes ou non a un thermocouple de temperature d'arret, sonde NACA, sonde"cobra", sonde "a bille" a 3 ou 5 trous, sondes a "3 doigts", a "4 doigts", sondes cylindriques, sondesa pression statique a orientation automatique ou manuelle. Les rScepteurs sont Sgalement tres varies :capteurs multicanaux, enregistreurs, lectures et photographies sur colonnes a eau ou a mercure, systemesSlectroniques avec transmission directe sur ordinateur, cartes, bandes magnetiques, etc...

Les valeurs Stant connues en chaque point, certains participants conseillent d'utiliser les rSsultats endSterminant I'Scoulement uniforme Squiyalent calcule apres melange en conservant le debit, la quantitede mouvement et 1'Snergie. Notons que pour mesurer directement la tempSrature, par exemple, dans I'Scou-lement homogene, les sondes doivent etre placees suffisamment en aval et 1'hypothese d'adiabaticite doitetre verifiee.

4.3. Moyens de mesure a reponse rapide

Pour clore la discussion, une utilisation intensive de mesure a reponse rapide est vivement souhaiteeet meme dSja lancSe. En effet, tous les constructeurs presents reconnaissent qu'il existe un Scart entrela valeur moyenne dans le temps de la mesure obtenue par 1'amortissement naturel d'une ligne a rSponselente et la moyenne des mesures directes effectuSes par des capteurs a rSponse rapide. De plus, dansles Scoulements avec ondes de choc 1'Scart est plus sensible.

Chaque systeme de mesure a ses caractSristiques particulieres. Fils chauds, films chauds et capteursa jauges montes sur pivot sont utilisSs pour la dStermination des sillages de rotor, systemes piezo-Slectriques ou a rSluctance,strain gages pour les variations de distribution de pression.

Les besoins et les recommandations les plus cites sont : miniaturisation, facilitS et stabilitS de 1'Sta-lonnage, tenue mScanique, facilitSs de montage, tenue des performances jusqu'a des tempSratures de 350°ou plus.

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TESTS OF COMPLETE COMPRESSOR OR TURBINE

REYNOLDS NUMBER EFFECTS

COLD TESTING OF TURBINES

27

TESTS OF COMPLETE COMPRESSOR OR TURBINE

Determination of stage characteristics

All contributors on this subject measured the static pressure between blade rows with flush outer walltappings. This provides insufficient data to derive stage characteristics and so other assumptions haveto be made about flow conditions. The usual method is to assume the design flow angles and either thedesign distribution of total temperature or a stage efficiency equal to the measured overall polytropicvalue. ROLLS ROYCE, DERBY assume that each blade row exhibits an off-design behaviour pattern similar tothat of. a controlled velocity ratio cascade test.

Stator mounted total pressure and total temperature probes are being increasingly used in both the UK andUSA. The results from these probes are fed into a streamline curvature type of through—flow computerprogram to build up a picture of the flow pattern in the turbomachine. A few assumptions are still requi-red but this approach can give some ideas about blade element performance as well as the normal stagecharacteristics. The best configuration for a stator mounted probe has not yet been determined. One Ame-rican manufacturer removes the stator completely and replaces it with an instrument rake whilst a Britishfirm has found that it is best to mount the probes back from the stator leading edge to avoid them beingdamaged. Another problem is that the presence of the extra instrumentation can significantly alter theperformance of the turbomachine (one American multistage compressor lost 1 to 2 per cent efficiency).

Some research workers had made traverses between the blade rows in multistage compressors but all expres-sed doubts about the absolute accuracy of their results. The traverse probes were always too large (formechanical reasons) compared with the blade pitch and the axial gap between adjacent blade rows. The probeblockage affects the local rotor performance and the static pressure field around the stators. Howeverwith suitable adjustments, traversing between blade rows produces the best reconstruction of turbomachineperformance with the minimum of additional assumptions. Also it provides some insight into wall boundarylayer growth and other end effects.

Very few turbomachine tests appear to have been done where sufficient reliable data were recorded so thatthe complete flow patters could be deduced without any extra assumptions. One research machine that wasfully instrumented was the NACA five-stage compressor reported in NACA RM E-56 G-24.

Stage characteristics are not usually derived for multistage turbines but inter-blade row measurementsare made for comparison with design values.

Isolated stage characteristics are not usually available for stages tested in multistage compressors. Ifthey were, they would probably differ from the derived multistage characteristics because'of differentupstream and downstream conditions and different range of operation. American experience has been thatsingle stage characteristics can be used to predict the overall performance of an axial-centrifugalcompressor. „

Determination of shaft power

Most people determine the power output of a turbine by a water brake or a swung brake of some type. Thecorrections for these brakes can be accurately determined, so a direct measurement of shaft power issuperior to the alternative methods, recording mass flow and temperature drop.

The compressor situation on power input measurement is far more undecided. Various types of torquemeterare used and all require empirical corrections most of which are quite significant. It appears that ifthe corrections are accurately determined and if the torquemeter is installed and used with care, thenit can produce good results. Direct measurement of temperature rise is simpler and produces consistentresults although the absolute values may not be as accurate as from a torquemeter when used in the bestway. N.G.T.E. did some comparisons with a large two-stage fan and found that the torquemeter and tempe-rature rise efficiencies agreed to within * 1 per cent. An American firm quoted a similar comparison fora small centrifugal compressor where a careful accounting of parasitic losses and heat transfer resultedin an isentropic efficiency measured by the torquemeter agreeing with temperature measurement to within2 per cent. However most of the American industry prefer to base their efficiencies on direct measurementof temperature rise. European practice is fairly evenly divided between the two methods and one firm usesa torquemeter for single stage tests and temperature measurement for multistage compressors.

The most commonly used torquemeters are optical, phase shift or variable inductance types. Mechanical andwindage losses have to be determined for all torquemeter applications. In large machines these parasiticlosses may be relatively small and so high accuracy is not necessary. However they must be determined moreprecisely for low flow turbomachines where the parasitic loss can be a significant proportion of the powerbeing measured. The mechanical losses are associated with rig and compressor bearings, couplings, seals,gears, etc.. whilst the windage losses may account for leakage as well as the aerodynamic drag on therotor drum and other surfaces. For some torquemeter installations a further significant correction arisesfrom the temperature effect on the torque shaft rigidity.

Automatic data logging

Most firms and research establishments now use automatic data loggers, coupled with pressure transducers,to record test measurements. As with torquemeters, it seems that a fair amount of care and attention isneeded to produce good results and also an initial development period is required for most data logger

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systems. Lack of complete confidence in the automatic system is apparent because the more important flowparameters are usually duplicated on manual instrumentation. Also visual display of a few parameters isoften necessary to set the test conditions.

If suitable range pressure transducers are used and correct calibrations obtained then data loggers areas accurate, or perhaps slightly better, than a direct manometer reading. However a significant gain inoverall accuracy arises because a complete set of readings can be taken within a few seconds comparedwith the minutes required to read several banks of manometers. Other advantages of automatic systems arethe possibility of numerous data points in critical regions (e.g. near surge), a saving in overall testtime and simplification of the analysis process.

Measurements in the rotor

Alternating stresses are commonly measured on rotating blade rows both for mechanical safety and also toinvestigate vibrations phenomena. Strain gauges are attached to the blade surface and the analogue signalis conveyed to the static frame either by slip rings or telemetry. These latter two methods are also usedto transmit thermocouple signals when measuring blade surface temperatures.

Several research organisations have successfully measured the static pressure distribution around thesurface of a rotating blade. In most cases this work has been limited to low speed rigs but LIVERPOOLUniversity have succeeded in measuring static pressures in a centrifugal impeller rotating at 8000 rev/min. The impeller is about 500 mm diameter and contains 64 pressure tappings which are piped to a scanningvalve system on the axis; a single plenum chamber is then used to transmit the pressures from the rota-ting system to the stationary instrumentation. There have been mechanical difficulties with the systembut a more important problem concerns the correction for the pressure difference due to rotation betweenthe static tapping and the axis of the machine. Since the pressure lines run along the surface of therotor remote from the impeller passages, the air temperature and hence density is difficult to eithermeasure or assess. Another significant piece of research is being done by DFVLR at PORZ-WAHN. They aredeveloping a la.ser-doppler velocimeter which might allow them to measure the local flow velocities withinrotor blade channels.

Blade tip measurements

For low speed rigs it is satisfactory to measure the static tip clearance with feeler gauges and assumethat the running clearance is the same. The static clearance is also measured on high speed compressorsand turbines but the operating clearance has to be determined separately. This is usually done withinductive or capacitive type proximity probes. Abradable probes indicate minimum tip clearance and cut-wire type instruments are sometimes used as safety devices.

If the tip clearance is very small then a change may have little effect but above a certain level ofclearance rotor performance may degrade rapidly. For a compressor the usual effect of increasing rotortip clearance is to reduce mass flow, efficiency and surge margin. ROLLS ROYCE suggest that a 1 per centincrease in tip clearance as a proportion of the stage inlet area produces a 5 per cent loss of thestage peak pressure rise. CAMBRIDGE University are attempting to predict the effect of tip clearancewith a three-dimensional boundary layer theory. Increasing the rotor clearance in a turbine also reducesthe efficiency but it may increase the mass flow. N.G.T.E. suggest that an increase of 0.1 per cent inthe clearance/blade height ratio produces about 0.15 per cent loss of efficiency.

High frequency pressure transducers mounted in the outer casing can be used to locate tip shock patternsin transonic compressor rotors. The results must be interpreted considering the physical size of thetransducer compared to the blade and shock thickness. American experience has shown that the transducerrecordings vary considerably from passage to passage and, although any results on shock strength may notbe meaningful, useful information on shock position can be obtained. O.N.E.R.A. have compared shock posi-tion measurements in a rotor and in cascade and in both cases they discovered that the static pressuremeasurements gave a shock location ahead of the Schlieren-indicated position. It appears that the pres-sure rise caused by the shock travels upstream in the wall boundary layer. DFVLR suggested that thiseffect is small below inlet relative Mach numbers of about 1.2.

At the tip of a centrifugal impeller the flow conditions are unsteady, the Mach number is usually highand the flow passage very narrow. In these circumstances, detailed flow measurements obtained by traver-sing with miniature probes are of doubtful accuracy. However a knowledge of the flow pattern in thisregion is necessary and so traverses have to be made even if they only produce qualitative results.

Measurement of unsteady flows

Oscillating pressures are usually measured with piezo-electric, capacitive or semiconductor type transdu-cers. These devices have a high resonant frequency (above 100 kHz) and the latter type can be made verysmall so that they can be built into pitot probes. The disadvantage of these transducers is their depen-dence on operating temperature which limits their application to dynamic pressure measurements.

Hot wire and hot film anemometers are commonly used to obtain continuous readings of flow angle, velocityand temperature in a turbulent stream. Unfortunately these instruments are very fragile and alternativesystems are under investigation. In America a number of flow velocity and flow direction probes utilizingminiature high frequency pressure transducers have been tried and may be useful in certain frequency ranges.

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The work at DFVLR" on a laser-doppler technique has already been mentioned. It is at present limited tovelocities below 50 m/s because of problems with the electronic system.

The determination of correct time-average values of flow parameters in a strongly fluctuating flow isvery difficult. Inert measuring instruments such as a pitot tube and manometer tend to produce valueswhich are greater than the true time average. The error depends on the amplitude and form of the pressureoscillation and on the geometry of the measuring system. DFVLR have developed three different methods ofobtaining correct time average pressures :

(1) a mechanical device with a rotating piston(2) a hydraulic device with fluid filled capillary tubes(3) a special pneumatic system in combination with a pressure pick up and a computational method of

correlation.

The Americans have used electrical averaging of the analogue signal in cases where the response of themeasuring system is well above the frequency of the parameter being recorded. In other cases the sensingelement has been deliberately slugged to match its response to that of the data recording system.

REYNOLDS NUMBER EFFECTS

For axial turbomachines the effective Reynolds number is usually defined by :p V cReynolds number =

where :

p is the static density

V is the relative velocity at inlet for compressors and at outlet for turbines

c is the blade chord at mean radius

U is the viscosity

A similar formulation is used for radial machines but in this case the relevant dimension is either thetip diameter or the hydraulic radius. The Americans believe that a general Reynolds number correlationdoes not exist because of the strong effects of stage matching and other associated parameters such asturbulence level. In view of this, they prefer to consider performance changes on a particular machineusing the parameter

(T477)(Ti + 198'7)Reynolds number index :T 2

Ojjl) x 717.4

where :

P = inlet total pressure

T = inlet total temperature

198.7 and 717.4 are constants from the Sutherland viscosity expression.

Transition and separation of the blade boundary layer is influenced by several other factors apart fromReynolds number. Flow turbulence, surface roughness and the pressure gradient all influence the boundarylayer conditions. Some systematic cascade experiments at LIVERPOOL University showed that Reynolds numberand turbulence level are Independent parameters. There are problems in defining the true turbulence levelat entry to a turbine and throughout a multistage compressor.

Turbomachine components are usually tested either at engine size or at thelargest scale that can beaccommodated in the particular test rig. This choice is dominated by manufacturing considerations, ins-trumentation requirements and the need to evaluate as nearly as possible the engine size aero-mechanicalproblems. In order to minimise Reynolds number effects it is desirable to achieve component test valuesgreater than about 2 x 10^. Above this critical value, quite a wide change of scale appears to be accep-table. For basic exploratory development of aerodynamic concepts, large low speed or water rigs have beenfound to be an effective tool. Such rigs are useful because Reynolds number effects are usually of secon-dary importance to choice of stage parameters in establishing turbomachinery performance.

Experimental determination of Reynolds number usually covers a relatively small range. There are fewresults on supercharging at the machine inlet and most pressure change information has been,obtained byinlet throttling. In one extreme case, ROLLS ROYCE covered the range of Reynolds number from 105 to 106but a range of between 2/1 and 3/1 is more normal. Most Reynolds number tests of this type are done withone compressor or turbine build, ie with constant blade tip clearance.

The correct relative clearance may be used when Reynolds number tests are done with machines of differentscale but again only a narrow range of Reynolds number has been examined. Blade surface roughness is ano-ther relevant parameter which has been ignored in mbst experimental work.

Changes of Reynolds number above a critical value of about 2 x 10 appear to have little or no effect onthe performance of a turbomachine. Below this critical value most machines lose efficiency with reducing

30

Reynolds number; some compressors also lose mass flow and surge margin. Because of variations in stagematching, multistage compressors can respond to a change of Reynolds number in many different ways. Inorder to predict the effect of Reynolds number on axial compressor performance, several firms use theROLLS ROYCE method presented in the Journal of Engineering for Power, Volume 90, April 1968. A recentReynolds number correlation for axial turbines has been given by Craig and Cox in the Proceedings of theInstitution of Mechanical Engineers, Volume 185, 32/71.

COLD TESTING OF TURBINES

Aerodynamic research and development of turbines is mainly done with rigs using cold air. The advantagesof cold rather than hot tests are :

(1) mechanical design of the turbine and rig is easier(2) cheaper materials can be used(3) instrumentation is simpler and likely to be more accurate(4) a cold air supply is easier to produce(5) all the previous features reduce cost

Hot testing of turbines is primarily done for heat transfer research but hot tests may also be done fordurability testing or to verify the performance of turbines developed under cold flow conditions.

The aerodynamic effect of the coolant discharge is simulated during cold turbine tests by maintaining thecorrect coolant/mainstream mass flow and density ratios. Alternatively, other coolant/mainstream flowparameters such as momentum ratio can be set because the difference between the various methods has onlya second order effect on turbine efficiency.

The total pressure distortion and flow turbulence created by a combustion chamber are likely to have asignificant effect on the heat transfer characteristics of a turbine but probably only have a small effecton the aerodynamic performance. Hot turbine tests are invariably done behind a combustor and thus correctsimulation is achieved. Some ROLLS ROYCE cold turbine tests where inlet turbulence was simulated firstwith a gauze and then with large cylinders produced no significant effect on turbine performance. Howeverthere could be an effect due to inlet maldistribution. Combustion chamber delivery conditions are largelyunknown but if they were, accurate simulation of both turbulence and pressure distribution could be verydifficult.

The blade profiles used in the cold air tests are normally retained for the engine turbine. The effect ofthe different specific heat ratio is considered to be small and of the same order as other rig to enginechanges (eg tip clearance, seal leakage, disc windage). An empirical correction (of the order of 1 percent on efficiency) is made by some firms to allow for all these differences.

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COMPRESSOR STABILITY AND DISTORTION TESTS

33

COMPRESSOR STABILITY AND DISTORTION TESTS

Question 1 : What parameters do you use to express steady and unsteady distortion numerically and how doyou measure it ?

It was evident in the discussion of this question that the answer itself depended to a large extent uponthe purpose for which the distortion factors were to be used. Clearly, the use of distortion factors hasits justification in the simplification offered through their use. Thus, a successful factor correlatesthe behavior of a large number of separate distortion patterns and reduces the requirements for extensivetesting of prospective inlet and engine combinations. A successful distortion factor also allows the useof artificial devices (usually screens) to simulate the expected patterns to be encountered in operationwith the eventual inlet and, of course, allows the inlet behavior to be described in terms of the valuesof distortion parameter produced.

A possible future development of distortion factor usage would allow the replacement of highly turbulentflows with "equivalent" flows utilizing a "quasi-steady" distortion factor. Such a successful quasi-steadydistortion factor would again allow utilization of such straightforward and relatively inexpensive testtechniques as the use of screens. These tests would be in contrast to tests utilizing high levels of tur-bulence, which have the attendant requirements of expensive instrumentation and enormous data reductiondemands.

The viewpoint was also strongly expressed, and generally agreed to, that the results of tests utilizingdistortion factor concepts should be regarded as a convenience, primarily of use in the earlier stages ofinlet and engine development, and that every effort should be made to test the engine and inlet combina-tion as early as possible in a development program.

Generally speaking, all attempts to describe distortion involve the description of the amount of azimuthal(9) distortion and the amount of radial distortion. Usually separate factors are utilized for each ofthese forms of distortion, but in some cases a single factor, formed as a combination of the radial andazimuthal distortion factors has been utilized. In every case discussed the prime emphasis was placed onthe description of the azimuthal distortion factor for the rather obvious reason that if for some reasonthe radial distortion encountered is unavoidable, then it can be designed for, whereas this is quiteobviously not possible in the case of azimuthal distortion. Some further consequences of this situationare discussed below.

It was evident that all azimuthal distortion factors in general use today, incorporate in their descrip-tion, in some way, the well known frequency response of an airfoil to a fluctuating disturbance. The most.commonly utilized factor throughout Europe is the DC-60, which in essence integrates the (non-dimensional)stagnation pressure decrement in a sixty degree segment.(In practice, the critical segment width is inves-tigated and is not necessarily 60°). The largest such value on the circumference forms "DC.60". It isimplied, of course, that the airfoils have little response to stagnation pressure decrements of significan-tly less span than 60*, and also that not significantly greater response is experienced for pressure decre-ments of ereater extent than 60°, the airfoil bv this time behaving as if in quasi-steady state.

An alternative method of incorporating the same fundamental aerodynamics is utilized in such factors asK (Pratt and Whitney). In this class of factor the azimuthal distortion is fitted by a Fourier series,and a weighting function reflecting the decreased airfoil response to the higher frequencies, is appliedto each of the coefficients.

It is natural to hope that a "universal" distortion factor, reflecting the best characteristics of themyriad of distortion factors now available for particular engine types, could be developed. It is in thiscontext, however, that the preceding remarks concerning the radial distortion bear repeating. Thus, giventhat the radial distortion has at least some significant effect, it is precisely the fact that such dis-tortion "can be designed for" that disqualifies the concept that a generally valid distortion factor inclu-ding radial effects could be formulated. This statement is not necessarily true for "purely azimuthal"distortion, of course, and the establishment of a widely applicable azimuthal distortion parameter wouldbe of great utility to the industry. It was the prevalent feeling in the meeting that at some time in thedevelopment process of a given flight vehicle the radial distortion should become of relatively littlesignificance, which further justifies the focusing of effort being directed to the azimuthal distortionfactor.

A further use of distortion patterns appears when two rotating components are to be matched. Thus, forexample, if a fan-low pressure compressor combination is to be matched to a high pressure compressor, itis important to have a knowledge of how inlet distortion into the fan-low pressure compressor combinationis attenuated prior to entry into the high pressure compressor, it is important to have a knowledge ofhow inlet distortion into the fan-low pressure compressor combination is attenuated prior to entry intothe high pressure compressor. Here the successful utilization of distortion factors would allow the simu-lation of the outlet distortion of the one rig at the inlet of the other. Here again, it is apparent thatit is desirable to test the complete system as early as possible in a development program.

The characterisation of distortion when large time dependent fluctuations are present was also considered.It is current practice to characterize such time dependent distortion through the use of the same distor-tion patterns as already utilized in the description of steady state distortion. In much the same way aswere the parameters originally adjusted to account for the airfoil response characteristics, however,allowance must be made for the reduced effect of the high frequency portions of the time dependent dis-turbances.

Generally, two techniques of describing time dependent distortion are utilized. In the first, actual timehistories of either the distortion parameter itself, or time histories of stagnation pressures (or

34

temperatures) that make up the distortion parameter, are searched. By searching a statistically meaningfulquantity of data, the maximum value of distortion parameter expected, for the given operating condition,can be estimated. Provided that the distortion parameter used for steady state purposes was sufficientlyaccurate, it is then assumed that the "instantaneous" distortion factor obtained as described above willserve the same purpose. There are two principal problems with the use of instantaneous distortion factors.Firstly, the number of samples of the data required to provide a meaningful statistical sample is verylarge, and the data processing cost consequently large also. Secondly, without the additional informationof the frequency content of the distortion, there is no way to judge whether the observed peak in distor-tion would have been of sufficient duration to cause the full response appropriate to the distortion.

The second technique for describing time dependent distortion involves the use of statistical averages ofthe disturbances, and the construction of "pseudo-steady" distortion parameters from the statistical ave-rages. In its most direct form, for example, the root mean square (R.M.S.) value of stagnation pressurefluctuations at each probe would be multiplied by an appropriate (usually semi-empirical) factor and addedto the steady state value of the stagnation pressure existing at the probe. The resulting "deepened"pattern is then utilized to calculate the effective distortion pattern. This technique has the advantagethat less data processing is required than that for the instantaneous distortion parameter technique. Also,if desired, the higher frequency inputs can be electronically filtered with comparative ease. The majorproblem with the technique is in determimining what should be the appropriate semi-empirical factors to usein multiplying the root mean square values of the fluctuations.

It is possible to utilize more statistical information than simply the R.M.S. values of the probe readings,but of course, at increased complexity and expense. For example, it would seem apparent that valuable in-formation would be obtained by considering the correlation and coherence of the various signals. By sodoing, estimates could be made of the circumferential extent of the various timewise disturbances, infor-mation which would be of direct use in estimating the airfoil response to the disturbance. The opinion wasadvanced that such techniques tend to be prohibitively expensive to justify the slight increase in predic-tive capability that they might afford over the use of the straight R.M.S. values for design purposes. Itwas felt also, however, that coherence and correlation studies were of great use in research areas wherethe fluid mechanics of actual surge inducing events, etc... were to be studied. Such techniques are alsoparticularly important for investigating disturbances created by system resonances.

Regarding the use of distortion factors, it was pointed out that if a very large quantity of data was to bestudied, a factor involving a linear combination of the pressure signals would lead to simple electroniccircuits for on-line processing of the distortion factor. This, of course, is the case for factors of thetype of K , described earlier. A comment was made to the effect that we should guard against adapting toosimple a factor just to aid calculation.

A final point was raised concerning the number of probes required for successful description of a distor-tion pattern. The commonly accepted number for use on full scale tests is presently forty probes and therewas some question as to whether such a large number of probes was required. Several studies were mentio-ned in which, mathematically, successive probes in a pattern were removed and the resulting pattern pre-dicted, recalculated. It appeared that in most cases little significant change in the predicted distortionparameter occurred till many (up to fifteen) probes were so removed, suggesting perhaps, that fewer probesmight suffice. A final comment was added that in most such tests some probe breakage occured, and someredundancy of measurement was useful.

Question 2 : How do you simulate steady or unsteady-distortion of pressure and temperature ?How do you simulate the turbulence level ?

Most investigators use screens or grids to simulate steady state distortions, and the use and limitationsof such devices is quite widely known. In the field of unsteady distortion, two classes of device tend tobe utilized. In the first class, very large scale turbulence is generated in an essentially random manner,but possibly with predetermined spacial variations in the average fluid properties. In the second class,discrete frequencies of disturbance are generated.

An example of the first class of unsteady distortion generator is the so called choked plug inlet, inwhich a convergent-divergent duct is operated in a supercritical condition. This condition leads to strongshock waves and strong wave-boundary layer interactions which in turn produce very high turbulence levels.The turbulence level may be varied by axially translating the "plug" so as to change the degree of criti-cality, and the spatial variation of turbulence intensity may be varied by shifting the plug to a non-axi-symmetric location.

Several examples of the discrete frequency class of "turbulence generator" have been constructed, such asrotating spike devices operate by rotating segments of screens or perforated plates past similar stationarysegments, with the result that a discrete frequency of disturbance is produced.The pneumatic devices injectquantities of air in a pulsating manner, usually on radial arms. An alternate approach is to provide cir-cumferential slots in the outer casing of the inlet duct, and by rotating a slotted band around the circum-ferential slots, pulsate the static pressure and hence mass flow rate, in the duct.

The utility of the discrete frequency turbulence generators arises mainly in their application to theresearch aspects of turbomachinery. Thus, by locating frequencies to which the engine is particularly sen-sitive, considerable light can be shed upon the underlying fluid mechanics of the interaction system. Reso-nances can also be located by driving the disturbances at the resonant frequencies and observing the systemresponse.

The random turbulence generators find their greatest utility when used as simulators of inlet generated tur-bulence. Their prime usefulness arises mainly in the early stages of engine development when the inlet is

35

not yet ready for matching to the engine. It was pointed out that even when a large wind tunnel is avai-lable for inlet and engine test, model tests are usually over by the time the inlet and engine are opera-ted together in the tunnel.

It is quite evident that the appropriate use of such devices is somewhat open to debate, what with thetypical "engine man" preferring that the relatively large expenses involved be directed more to inlet deve-lopment so as to reduce the levels of distortion encountered. Contrarily, the typical "airframe man" pre-fers that the engine be run wikh very large surge margin so as to be insensitive to virtually any distor-tion provided. In such a debate the turbulence generator appears as the device which will hopefully allowengine development to the stage where the engine will accept a given level of turbulence with minimumpenalty in surge margin. No real consensus was reached on these varying viewpoints.

More classically, investigators have been concerned with the effects of homogeneous turbulence upon engineperformance. Grids have traditionally been used to generate such turbulence, and experiments were describedin which high density screens were utilized to generate levels of turbulence up to 7 % at a.Mach number ofabout 0.4. It was found that large scale instabilities in the flow field occurred if the open area ratiowas1 less than 0.56. This level of turbulence appears to be the limit attainable with such devices, higherlevels being attained only through the use of the previously described turbulence generators. In the lattercase, of course, the turbulence is usually far from homogeneous.

Question 3 : Do bench measurements of distortion and compressor behavior behind such simulators agree withflight observations ? (Specify the flight conditions where the agreement is satisfactory).

It was generally agreed that satisfactory agreement has been obtained in most flight regimes at subsonicMach numbers. The agreement in these cases involves the comparison of bench tests utilizing screens, etc..for simulation of steady state distortion, and the actual flight measurements.

The agreement for operation at supersonic Mach numbers, or at some particularly severe operating conditions(high angle of attack, etc..) was found to be less satisfactory. It is apparent, of course, that theseflight regimes correspond to conditions where turbulence effects are expected to be large. It was statedthat the use of turbulence generators had improved the agreement over that found with screens only, butthe interpretation of the results has been marred, to date, by the difficulty of obtaining adequate flightdata in these regimes. The major problem has been that of obtaining appropriate flight instrumentation.

Question 4 : How do you define and measure surge points ?

Many techniques were mentioned for measuring surge points, such as separately or in combination : hotwires, strain gauges, static pressure pick-ups, stagnation pressure pick-ups and ears. The quantitiesmeasured (separately or in combination) were outlet pressure, flow rate, inlet pressure, pressure ins-tability and noise.

Generally in "high energy tests" (high pressure ratio tests) a clear distinction exists between stalland surge. When the compressor (or fan) is operated at decreasing mass flow along a constant speed line,a stall is encountered prior to the onset (with further decrease in mass flow) of surge. These regimescan be confused if inadequate instrumentation is provided. It was the feeling of the meeting that therewas little doubt when the operating limit of the compressor or fan was approached, and consequently, noreal reason exists to attempt to precisely determine a quantitative definition of the onset of stall orsurge.

When the compressor or fan is operated at very low pressure ratios, the separation of stall and surge isnot always possible. Tests were mentioned in which a compressor with strong tip spoiling was operatedwell into the stall regime, but surge did not occur.

Question 5 : What effect do the inlet and exhaust ducting have on surge points ?

Here again, as in question 4, the behavior of the device depends to a large extent on the operating range.When high pressure fans or compressors were studied, it was the experience of all investigators that theinlet and exhaust ducting had only very slight effect upon the location of the surge points. Quite obvious-ly the behavior during surging was dependent on the ducting, primarily through the effect of the volume ofthe ducting upon the surge frequency.

When low pressure ratio fans or compressors (particularly such low pressure devices as ventilation fans)are considered, the location of their surge points is found to depend, to a high degree, upon the geometryof the inlet and exhaust ducting as well as upon the condition of the casing boundary layers upon entry tothe test device. Several studies in the open literature have considered these interaction effects of surgepoint behavior and ducting volume.

Question 6 : What technique do you use for investigating rotating stall ?

All investigators used singly or in combination, hot wire anemometry, high frequency response pressure ins-trumentation or strain gauges. It was emphasized that the use of multiple probes was necessary to distin-guish the number of stall cells occurring azitnuthaly. It was also suggested that the use of several probesin the radial direction would be desirable for research purposes.

37

QUESTIONS D'INTERET GENERAL

39

L'un des premiers points discutes au cours de la reunion consacree aux "questions d'intere't general"fut de tenter de quantifier 1*importance des diverses motivations qui incitent a entreprendre des essais surelements de turbomaohines. II a ete reconnu que la premiere raison de 1'importance de ces essais est l'inter§ttechnique des informations fournies. Ceci n'est que la confirmation des impressions degagees au coura des dis-cussions precedentes consacrees a 1'interpretation et a la valeur des essais partiels. Ceci etant admis, ilexiste deux incitations : I'une fondee sur des considerations d'ordre financier : il est evident que les essaissur elements partiels cofttent moins cher que sur turbonachine complete; 1'autre ineitation est basee sur legain de temps. II n'a pas ete possible de dormer une importance plus grande a I'une ou 1'autre de ces incita-tions. Ceei provient peut Stre en partie du fait qu'elles sont Intimement liees, car si le coflt d'un travailest faible, la raison en est la plupart du temps que ce travail est de faible dur e.

Les participants ont discute egalement du but poursuivi par les essais partiels, "Recherche etameliorations des connaissances" ou "Developpement des turbomachines". Bien que la reunion groupait aussi biendes Industriels que des Organismes de recherche, les reponses n'ont pas donne une separation nette. Tous lesIndustriels reeonnaissent bien une priorite au but "Developpement des turbomachines" mais admettent que 1'ob-jeotif "Recherche" represente une valeur d'environ 30 /£. Cependant les essais de "recherche" sont en prinoipebien integres dans le prooessus eonduisant a la mise au point d'une turbonachine. D'une maniere generale lesessais effectues par les Industriels peuvent se classer en trois chapitres i

a) Recherches de base : elles rassemblent tous les essais susceptibles de dormer des informations pourfatre un projet valable (par exemple essais de grilles d'aubes),

b) Recherches d'idees nouvelles : il ne suffit pas pour faire un bon moteur de s'inspirer d'un modeleancien, il faut chercher a faire des progres. C'est pourquoi on entreprend des essais pour valider dessolutions nouvelles. Us necessitent 1'utilisation de nombreux resultats d'essais de recherches de baseet intervlennent directement dans 1'amelioration du moteur,

c) Essais des composants i ils permettent de prevoir les performances du moteur complet tout en dormantplus de facilite pour effectuer des modifications.

Quant aux organ ismes d'etudes, ils ont reconnu que leurs travaux sont axes sur les recherches et envue de I1amelioration des connaissances mais cependant insistent pour faire ressortir que 1'objectif lointainde leurs reeherches est d'aider les Industriels dans la mise au moint des turbomachines.

Les avantages apportes par les essais sur les elements partiels sont donnes ci-dessous :

- possibilite de connaitre un domaine de fonctionnement plus etendu en particulier exploration des domai-nes dangereux et connaissance plus approfondie du domaine hors adaptation,

- decouplage des problemes, diminution des interferences de fonctionnement entre organes, neilleured'adaptation du cycle,

- amelioration des methodes de conception et de calcul,

- solution plus facile et plus rapide de problemes et difficultes apparus sur moteur.

Bien entendu les avantages dependent du type d'essai envisage.

On constate que si on fait un classement, par priorites, des moyens d'essais utilises on trouved'abord le bane d'essai compresseur et presque en fin de liste les essais de grilles.

Cela signifie tout 1'interSt que chacun apporte aux essais de composants de turbomachines mais celane veut pas dire que les essais de grilles sont negliges.

Certes, il est vrai qu'en subsonique, on dispose deja d'un grand nombre d'essais de grilles etqu'on possede une tres bonne connaissance des calculs dans ee domaine. Mais, si ces cormaissances permettent dereduire le nombre des essais de grilles, elles ne les evitent pas compietement.

D'autre part le developpement des recherches vers le haut subsonique fait qu'on ne peut plus utili-ser ce qui existe actuellement eomme resultats d'essais.

L'interSt des essais en grille en vue d'une application determinee reste done reel sauf peut-Streen supersonique ou 1"exploitation des resultats paralt aujourd'hui discutable.

Les participants sont d'acoord pour reconnaltre qu'on peut deduire les performances des turbomachi-nes gr&ce aux essais partiels. La precision est a priori et au mieux +_ 5 $• Mais au cours de la mise au pointd'une turbomachine le deroulement parallele des essais sur element et sur moteur complet donne une bienmeilleure fiabilite aux essais partiels de sorte que leurs resultats peuvent contribuer a 1'estimation desperformances a ± 1 %. Mais cette bonne precision necessite comme il vient d'Stre dit une eertaine connaissancedu moteur complet. En effet bon nombre de phenomenes non reproduits sur elements jouent un rOle important.Parmi ces phenomenes citons :

- 1'influence du nombre de Reynolds,

- jeu et fuites,

40

- non eonformite des champs aerodynamique s et plus generalement turbulence et distortions non reproduites,

- interaction entre etages,

- effets de la couche limite,

- ecoulement parasites,

- inertie thermique.

D'une maniere plus generale, les principales causes d1 incoherence, ou les "principftux traquenards"dans 1' exploitation des essais partiels sont dans 1'ordre :

1 ) Environnement ,

2) Champs de pressions,

3) Inerties thermiques.

Par mot "environnement" on entend les conditions physiques differentes sur le moteur de cellesexistant sur bane partiel. (Par exemple debit de fuite ou de ref roidissement , dilatations, jeux differents,etc..). II a ete reconnu que les debits secondaires, les couches limites ou les ondes de choc ont unstres grandeinfluence sur le calcul d'une turbomachine.

Toutes les causes d'erreurs, d'imprecisions ou d' incoherences Imposent par consequent d'executerles essais sur composants partiels avec beaucoup de soin et des precautions particulieres. II est en particu-lier recoramande lors de ces essais :

- d'utiliser de preference les pieces des machines

- de reproduire si possible les prof ils des vitesses,

- de contr&ler les resultats en comparant les essais partiels et essais sur moteur. En particulier iln'existe aucune methode autre qu'empirique d'etablir une correlation entre les resultats des grilles etresultats d'etages tournants,

- de faire des tests de recoupement interne (recoupement par le debit, contrOle des temperatures, etc..),

- de traiter les mesures par des methodes statistiques,

- d'avoir une instrumentation nombreuse.de bonne qualite et bien adaptee au probleme etudie.

Le fait que I'aspect "instrumentation " ait ete mis en fin de liste ne signifie point que ce pointest moins important. Bien au contraire tous les participants ont exprime des besoins urgents et imperatifs enmatiere d1 instrumentation de mesures. Les caracteristiques les plus necessaires sont la miniaturisation et lestemps de reponse rapides. L'unanimite s'est degagee pour temoigner de besoins de mesures aujourd'hui difficiles(couplemetre) ou impossibles telles que vitesse et pression en mouvement relatif (rotation), ainsi que lesmesures des ecoulements instationnaires.

D'une maniere plus generale les participants pensent que des nouveaux moyens de mesures sont haute-ment souhaitables par exemple des interferometres a. laser, la teletransmission, la radiographie , les anemometrestree rapides, etc...

41

A N N E X

LIST OF COMPANIES AND INDIVIDUALS WHOPARTICIPATED IN THE STUDY

LISTE DES SOCIETES ET DES PERSONNES AYANTPARTICIPE A L'ETUDE

43

ALLEMAGNE:

- Deutsche Forschungs- und Versuchsanstalt ftlr Luft- und Raumfahrt e.V.Braunschweig (Institut fUr Aerodynamik)GBttingen (Aerodynamische Versuchsanstalt)Porz-Wahn (Institut fUr Luftstrahlantriebe)

- Institut fUr Strahlantriebe und Turboarbeitsmaschinen - Aachen

- Institut fUr Turbo-Flugtriebwerke - Stuttgart

- KIBckner-Huraboldt-Oeutz AG - Oberursel/Taunus

- Lehrstuhl fUr Flugantriebe - Darmstadt

- Lehrstuhl und Institut ftlr Flugantriebe - MUnchen

- Motoren-und-Turbinen Union GmbH - MUnchen

BELG_IQUE;

- ACEC (Gent et Charleroi)

- Fabrique Nationale Herstal

- Faculte* 'Polytechnique de Mons

- Institut von Karman - Rhode-St-GenSse

- University Libre de Bruxelles

ETATS-UNIS

- Airesearch Manufacturing Company - Phoenix

- AVCO Lyconing Division - Stratford

- Detroit Diesel Allison - Indianapolis

- Dynanalysis of Princeton

- General Electric Company - Cincinnati

- Iowa State University

- Lewis Research Center, NASA - Cleveland

- Pratt & Whitney Aircraft - East Hartford

- Teledyne CAE - Toledo

- University of Washington - Seattle

- US Army Eustis Dir AMRDL - Fort Eustis

- Wright-Patterson Air Force Base - Dayton

FRANCE:

- Centre d'Essais des Propulseurs - Saclay

' - CIT-Alcatel - BruySres-la-Chatel

- Ecole Centrale Lyonnaise - Ecully

- Microturbo - Toulouse

- ONERA - Chatilion

- SNECMA - Villaroche

- Turboraeca - Bordes

44

- Cambridge University

- Kilbride University

- National Engineering Laboratory - East Kilbride, Scotland

- National Gas Turbine Establishment - Pyestock, Farnborough

- Rolls-Royce (1971) Ltd - Bristol

- Rolls-Royce (1971) Ltd - Derby

- University of Liverpool

PAYS-BAS:

- Delft University of Technology - Delft

- Koninklljke Machinefabriek Stork NV - Hengelo

- Organisation for Applied Scientific Research - Delft

- Technological University - Eindhoven

45

PANEL STUDY GROUP MEMBERS

PIANKO, I.C.A. M. (Study Leader) STAe, Paris

CHAUVIN, Prof.J.

CHEVALIER, J.F.

DUNHAM, Dr J.

KUHL, Prof.H.

STAVEREN, Dr P.van

VKI, Rhode-St-Genese

SNECMA Villaroche, Moissy-Cramayel

NGTE, Pyestock, Farnborough, Hants.

DFVLR, Institut fur Luftstrahlantriebe, Porz-Wahn,Linder Hohe

Projectgroep Stromingsmachines, Institute for AppliedResearch TNO, Schoemakerstraat 97, Postbus 406, Delft

PANEL MEMBERS ALSO ATTENDING

DUCARME, Prof.J.

LANE, J.

MONNOT, G.

PAVLIDIS, Brig.Gen.A.

SEZGEN, Prof.

SURUGUE, J.

WINTERFELD, Dr Ing.G.

Universite de Lie'ge, Inst. de Mecanique

Rolls-Royce (1971) Ltd, Bristol

DRME, Paris

State Aircraft Factory, Athens

Middle East Technical University, Ankara

ONERA, Chatillon-sous-Bagneux


AUTHORS AND CONTRIBUTORS

BRAEMBUSSCHE, R.van den VKI, Rhode-St-Genese

GELENNE, P.

KIRSCH, Dr Ph.

MUYNCK, Rudi de

SIEVERDING, Prof.C.

BOUILLET, Roger

BURELLER, Ing.M.

CHAMPENOIS, Ing.M.

DEHONDT, Ing.

FABRI, J.

FAURY, Ing.General Marc

PLOT, Regis

FRIBERG, Jean

GAY, Bernard

Fabrique Nationale, Hestal

ACEC Charleroi

ACEC - Dock 52, 9000 Ghent

VKI, Rhode-St-Genese

SNECMA Melun-Villaroche

MICROTURBO-IDA, Toulouse

Centre d'Essais des Propulseurs de Saclay, Orsay

TURBOMECA, Bordes



Ecole Centrale de Lyon, Lab. de Mecanique desFluides, Ecully

CIT-ALCATEL, Bruyeres-le-Chatel


FRANCE

BELGIUM

FRANCE

UK

GERMANY

NETHERLANDS

BELGIUM

UK

FRANCE

GREECE

TURKEY

FRANCE

GERMANY

BELGIUM

BELGIUM

BELGIUM

BELGIUM

BELGIUM

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

46

AUTHORS AND CONTRIBUTORS (continued)

GIRAUD, Ing. en chef

MARTINAT, P.

MERIGOUX, Ing.J-M.

MISCHEL, Charles

PAULON, Jacques

ROUANET, Ing. Yves

SPETTEL, Florent

STOS, Pierre

THIAVILLE, J-M.

FOTTNER, Dr Ing.Leonhard

KURZKE, Dipl.Ing.Joachim

STARKEN, Dr Ing.Hans

DALEN, Ir.G.Van

HOPWOOD, D.J.

MACDONALD, A.G.

GOSTELOW, Dr J.P.

OLDHAM, R.K.

NORBURY, Prof.J.F.

GATES, Prof.Gordon C.

SCHMIDT, Marvin F.

TURBOMECA, Bordes


CIT-ALCATEL, Centre Pierre Herreng, Bruyeres-le-Chatel

MICROTURBO, Toulouse-Minimes






MTU Miinchen GmbH, Munich

Technische Universitat MUnchen, Lehrstuhl furFlugantriebe, Munich


FDO Engineering Consultants (part of VMF/Stork-Werkspoor), Hengelo

Rolls-Royce (1971) Ltd, Derby

Rolls-Royce (1971) Ltd, BristolEngine Division, Bristol

SRC Turbomachinery Laboratory, Cambridge

NGTE, Pyestock, Farnborough, Hants.

University of Liverpool, Mechanical EngineeringDepartment, Liverpool

University of Washington, Department of Aero & Astro,Seattle, Washington

USAF Aero Propulsion Laboratory, Wright-Patterson AFB,Ohio

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

FRANCE

GERMANY

GERMANY

GERMANY

NETHERLANDS

UK

UK

UK

UK

UK

USA

USA



P.T.O.

AGARD-AG-167621-135621-253621.001.4





P.T.O.

AGARD-AG-167621-135621-253621.001.4





P.T.O.

AGARD-AG-167621-135621-253621.001.4





P.T.O.

AGARD-AG-167621-135621-253621.001.4





as well as institutes or research organizations could take advantage of this oppor-tunity to compare their experiences, and their me'thods and make a fruitfulexchange of views. The present publication gives the synthesis and the generalconclusions of the study.





This AGARDograph comprises papers, discussions and conclusions from themeeting held at the Ecole Nationale Superieure de 1'Aeronautique et de 1'Espace,Toulouse, France, 18—21 September 1972.

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1972_AGARD-AG-167_Modern Method of Testing Rotating Components of Turbomachines

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