Objectives and approaches to research in turbine heat transfer are discussed. Generally, improvements in the method of determining the hot gas flow through the turbine passage is one area of concern, as is the cooling air flow inside the airfoil, and the methods of predicting the heat transfer rates on the hot gas side and on the coolant side of the airfoil. More specific areas of research are: (1) local hot gas recovery temperatures along the airfoil surfaces; (2) local airfoil wall temperature; (3) local hot gas side heat transfer coefficients on the airfoil surfaces; (4) local coolant side heat transfer coefficients inside the airfoils; (5) local hot gas flow velocities and secondary flows at real engine conditions; and (6) local delta strain range of the airfoil walls

Rohde, J. E.

English

NASA Technical Reports Server

TURBINE HEAT TRANSFER 
John E. Rohde 
National Aeronautics and Space Administration 
Lewis Research Center 
Cleveland, Ohio 44135 
Improved turbine durability and performance and reduced development cost 
will all result from improved methrids of predicting turbine metal 
temperatures. As you know, better metal temperature prediction methods 
require improvements in the method of determining the hot gas flo'''' through the 
turbine passage and the cooling air flow inside the airfoil and in the methods 
of predicting the heat transfer rates on the hot gas-side and on the 
coolant-side of the airfoil. The overall turbine heat transfer effort ~s 
directed at improving all four of these areas of concern. 
Achievement of these improvements requires a rigorous and systematic 
research effort from both the experimental and analytical sides. The 
experimental approach being pursued starts with fundamental experiments with 
simple shapes and flat plates; progresses on to more realistic cold, warm, or 
hot cascades; contiriues to progress on to more realistic warm turbine, large 
low-speed turbine, or transient turbine tests; and finally combines all the 
interactive effects in real-engine environment turbine tests. Analytical 
approaches being pursued also start with relatively simple mathematical models 
and progress to more realistic cases that include more interactive effects, 
and finally combines all the interactive effects of the turbine operating in 
the real engine environment. 
Cutrently, contract and grant activities are being (or will be) conducted 
to obtain fundamental experimental data and to develop and/or compare 
analytical methods Ln all four areas of concern. These contract and grant 
activities will be discussed in detail later in this meeting by the respective 
principal'investigators. 
Major NASA Lewis in-house turbine research efforts are being pursued to 
obtain more realistic .'lnd real-engine type turbine ~xperiments.' The NASA 
119 
https://ntrs.nasa.gov/search.jsp?R=19850002656 2020-03-20T21:16:31+00:00Z
Lewis Research Center is in the process of activating our High Pressure-High 
Temperature Facility (HPF) with initial 20 atmosphere and 25000 F 
experimentation scheduled for the last quarter of 1982. HPF will provide the 
country with a known real-engine environment in ,~hich to conduct controlled 
aerothermodynamic and structural research studies. We envision a multiple 
role for HPF in providing engineering-quality research data for modeling a9d 
code verification, in defining a real-engine environment, and in evaluating 
advanced turbine cooling technology in a real-engine environment. 
The major turoine research parameters of interest that will be measured or 
determined to provide a better understanding of the thermal, aerodynamic, and 
mechanical performance of air-cooled turbine airfoils are the following: 
1) local hot gas recovery temperatures along the airfoil surfaces, 
2) local airfoil wall temperature, 
3) local hot gas-side heat transfer coefficients on the airfoil 
surfaces, 
4) local coolant-side heat transfer coefficients inside the airfoils, 
5) local hot gas flow velocities and secondary flows at real-engine 
conditions, and 
6) local delta strain range of the airfoil walls. 
Currently, little of this type experimental research information exists with 
controlled warm or real-engine conditions and known boundary conditions. 
These in-house turbine research efforts will be conducted using the best 
available analyses to help define the test co~figurations, the types of 
research measurements, and/or the test conditions and for the compariso,n with 
the measured research'results. Analytical efforts will initially use the best 
available flow and heat transfer codes such as a two- or three-dimensional 
inviscid flow code and a two- or three-dimensional boundary layer heat 
120 
transfer code. These analyses will be applied at the mid-span section and 
possibly at the hub and tip sections or other local zones of the passage. 
More sophisticated three-dimensional V1SCOUS codes and three-dimensional 
viscous codes with boundary layer resolution will be used as they become 
available. These analytical efforts will be conducted using the best 
available source or sources in-house and on contract with industry and 
universities. 
121 
...... 
N 
N 
~:~~~~~~~~r~~~~and TURBINE ENGINE HOT SECTION TECHNOLOGV NI\SI\ 
Lewis Research Center 
TURBINE HEAT TRANSFER 
OBJECTIVES: IMPROVE ACCURACY OF PREDICTING LOCAL BLADE METAL TEMPERATURES USING 
COMPUTER CODES THAT ARE COMPATIBLE WITH STRUCTURAL ANALYSIS CODES 
APPROACH: • INVESTIGATE HOT GAS STREAM AND COOLANT PASSAGE FLOW MECHANICS AND 
HEAT TRANSFER 
• OBTAIN BENCHMARK-QUALITY AND ENGINEERING~QUALIrY DATA FOR EVALUATION 
AND IMPROVEMENT OF MODELS PRESENTLY USED IN PREDICTION CODES 
• UTILIZE IMPROVED MODELS TO IMPROVE ACCURACY OF PREDICTING LOCAL GAS-
SIDE AND COOLANT-SIDE HEAT TRANSFER COEFFICIENTS 
• UTILIZE FLOW MODELS TO IMPROVE ACCURACY OF PREDICTING LOCAL HOT GAS 
STREAM ENVIRONMEN~ THROUGH TURBINE ROWS AND COOLANT FLOW CONDITIONS 
INSIDE THE AIRFOIL 
• INTEGRATE IMPROVEMENTS IN PREDICTION OF HOT GAS STREAM ENVIRONMENT 
AND COOLANT FLOW AND LOCAL HOT GAS-SIDE AND COOLANT-SIDE HEAT TRANSFER 
COEFFICIENTS INTO IMPROVED METAL TEMPERATURE PREDICTION CODES 
• PROVIDE ENGINEERING-QUALITY TEST CASES FOR EVALUATION OF ACCURACY OF 
PREDICTION AN~ INPUT TO STRUCTURAL ANALYSIS CODES 
...... 
N 
W 
National Aeronautics and 
Space Adminislralion TURBINE ENGINE HOT SECTION TECHNOLOGV 
Lewis Research Center 
BUILDING BLOCK AEROTHERMAL TURBINE RESEARCH APPROACH 
/1 
..,,/ 
" / 
REAL 
WORLD 
NASA 
HIGH 
PRESSURE 
TURBINE 
ROTATION/i HIGH SPEED WARM TURBINE 
UNSTEADY I TRANSIENT TURBINE 
I 
CODE 
VER I FI CATI ON 
NI\S/\ 
~ ~~ 
-<) \,L. 
to.;. r~i)/ ~O~ Q:-~ ~~Q..Q.. LARGE 'LOW SPEED TURBINE CODE \"9: (' DEVELOPMENT ~~ ~~ 
/ 
/ 
AIRFOIL 
SPECIFIC I 
! 
I I 
STATIC, CASCADES 
COLD ANNULAR ANn WARM ANNULAR 
j FUNDAMENTAL I WIND ~UNNELS AND FUNDAMENTAL MODELS 
PHENOMENA I TRANSITION J VISUALIZATION J TURBULENCE J FILM COOLING 
/' ,/' 
. / 
/ /. / /. /' /' ,/ / ./ / ' , / /,
. //,.' . .. /' / ,/ 
\ 
\ 
... 
\ 
\ 
ANALYTIC ' 
MODELING 
...... 
N 
.;:. 
• II 
[ 
~~;:'and TURBINE ENGINE HOT SECTION TECHNOLOGY NJ\SJ\ 
Uooia_c..-
TURBINE HEAT TRANSFER 
PLANNING SCHEDULE 
ELEMENT FY I 81 L _~ _ J 83 I 84 I 85 I 85 I 87 I EXPECTED RESULTS I 
GAS-SIDE H,AI TRANSFER. NON-
ROTATING. ;l-D 
GAS-SIDE HEAT TRANSFER. NON-
ROTATING. FILM 
GAS FLOW ENVIRONMENT AND 
HEAT TRANSFER. NON-ROTATING 
GAS FLOW ENVIRONMENT AND 
~:aT TRANSFER. ROTAT.ING • 
MULTIPLE JET ARRAY 
IMPlfroEMENT 
COOLANT SIDE HEAT TRANSFER 
WITH ~OTATION AND ENTRANCE 
GEOMETRY 
~lETAL TEMPERATURE 
PREDICTION CODES 
IN-ltoU5E RESEARCH ANn 
VERI FI CATIONS 
,-
--DETERMINE INFLUENCE OF 
VARIABLES ON FLOW TRANSITION 
AND DURATION AND IMPROVED 
MODELS 
--SAME AS ABOVE WITH FILM 
COOL/N_G 
,----------.,------.,-------, --OBTAIN BENc'iiuARK QUALITY AERO-
'---_____ --1. __ - ___ ...J ______ ...I THERMODYNAMIC DATA Al'1D IMPROVED 
I~~-
l 
I _u 
THREE-DIMENSIONAL VISCOUS FLOW 
CODES. NO ROTATION 
1--------- ----:-, --SAME AS ABOVE WITH ROTATION I ______________ .J 
--IMPROVED HEAT TRANSFER CORRE-
LATION AND MODEL FOR IMPINGE-
MENT COOLING 
--HEAT TRANSFER CORRELATIONS. 
INCLUDING EFFECTS OF ROTA-
TIONS AND ENTRANCE GEOMETRY 
r - - - - - - - - - - - - -- - -, . --METAL TEMPERATURE PREDICTION 
L- -- - - - - - - -- - - ----, CODES WITH IMPROVED HEAT 
TRANSFER MODELS/CORRELATIONS 
--VERIFICATIONS OF FLOW. HEAT 
TRANSFER METAL TEMPERATURES. 
AND STRAIN PREDICTIONS, AT 
NEAR AND REAL-ENGINE TYPE 
CONDITIONS 
...... 
N 
1.11 
Nahonal Aeronautics and 
Space Administration TURBINE ENGINE HOT SECTION TECHNOLOGY I\JI\S/\ 
Lewis Research Center 
POSSIBLY CONTRACT AND/OR 
IN-HOUSE: 
SCOPE: 
DURATI ON: 
APPROACH: 
TURBINE HEAT TRANSFER 
METAL TEMPERATURE PREDICTION CODES 
REVIEW AND MODIFY EXISTING AIRFOIL METAL TEMPERATURE 
CODES FOR EFFICIENT INCORPORATION OF DEVELOPED 
MODELS AND FOR INTERFACING WITH STRUCTURAL ANALYSIS 
CODES 
THREE YEARS 
• REVIEW EXISTING CODES 
• INCORPORATE IMPROVED MODELS DEVELOPED UNDER HOST 
• ASSURE EFFICIENT INTERFACING WITH STRUCTURAL 
CODES 
..... 
N 
0'1 
National Aeronaullcs and 
Space Adminislration TURBINE ENGINE HOT SECTION TECHNOLOGY 
Lewis Research Center 
SCOPE: 
TURBINE HEAT TRANSFER 
IN-HOUSE RESEARCH AND VERIFICATIONS 
EXPERIMENTS AND ANALYSIS TO SUPPLEMENT CONTRACTUAL AND GRANT 
EFFORTS ON IMPROVING ACCURACY OF FLOW ENVIRONMENT AND HEAT 
TRANSFER PREDICTIONS AND THE VERIFICATION OF DEVELOPED/IMPROVED 
PREDICTION METHODS 
DURATION: SIX YEARS 
APPROACH: • MEASURE LOCAL HEAT TRANSFER COEFFICIENTS OVER A STATOR 
VANE AT NEAR-~EAL ENGINE CONDITIO~S AND COMPARE WITH 
PREDICT~ON 
• PROCURE LASER ANEMOMETER SYSTEM AND INSTALL AND CHECK-OUT 
IN NASA WARM TURBINE 
• OBTAIN MEASUREMENTS IN NASA HIGH PRESSURE TURBINE "(HPT) 
AT NEAR REAL ENGINE CONDITIONS TO EVALUATE PREDICTION 
ACCURACIES OF CODES FOR HOT GAS FLOW~ HOT GAS ENVIRONMENT~ 
HOT GAS-SIDE AND COOLANT-SIDE HEAT TRANSFER COEFFICIENTS~ 
METAL TEMPERATURES~ AND STRAIN 
NI\S/\ 
I--' 
N 
-...! 
ENGINEERING 
MEASUREfiENTS 
TO DEFINE 
'J/ 
TYPI CAL AI RFOI L AE ROTHE RMODYNAMI C 
ANALYSIS NEIAOD 
ENGINEERING HOT ·GAS 
BOUNDARY 
CONDITIONS 
..... 
~ 
TWO- OR THREE-DIMENSIONAL~ 
NON-VISCOUS OR VISCOUS 
FLOW CODE 
1< MEASUREMENTS TO 
T 
BOUNDARY LAYER HEAT 
TRANSFER CODE 
mGINEERING 7 
MEASUREMENTS TO ~ 
DEFINE AND VERIFY 
CODE 
\1/ 
ENGINEERING 
MEASUREr':ENTS TO 
VERIFY 1, 
VERIFY CODE 
LIFE NODE 
GENERATOR ........ "" 
THREE-DIMENSIONAL 
STEADY STATE AND 
TRANSIENT CONDUCTION CODE 
WALL ,--------, 
TEMP. STRUCTURAL 
AND ANALYSIS H PREDICTION 
PRESS. CODE CODE 
l' 
COOLANT-SIDE 
HEAT TRANSFER 
CORRELATIONS 
l' 
ONE-) TWO-~ OR THREE-
~ DIf1ENSIONAL VISCOUS 
FLOW CODE OR NETHORK 
CODE 
POSSIBLY ENGINEERING 
MEASUREMENTS TO MEASURE 
AND VERIFY 
. '~ 
ENGINEERING 
rEASUREMENTS TO 
VERIFY 
-N co 
• ,I •. 
National Aeronautics and 
Space Administration TURBINE ENGINE HOT SECTION TECHNOLOGY NI\S/\ 
Lewis Research Cenler 
ENGINEERING-TYPE MEASUREMENTS DESIRED 
T TI g,4, g,4 
P 9,4 
V~,4 
q/A -+-1\ 
rad -+-1 J 
-+-L./ 
~,Tw 
, 
, 
, 
T TI 
g,S, g,S 
I 
Vg,s 
I 
Vg,V 9 
-.... ql A tat 
~~II 
,-Tw 
~-, { !J ~ \'. ,- E 
'_ ... q/A
rad 
.- ON TIP 
SHROUD 
T,N 
• 
(~);- q/A d ON ENDWALL 
\ ra 
.... _/ 
P g,6 
...... 
N 
lO 
National Aeronautics and . 
Space Administration TURBINE ENGINE HOT SECTION TECHNOLOGV NI\SI\ . . . 
Lewis Research Conter 
NONINTRUSIVE·VELOCITY MEASUREMENTS , 
. THROUGH COMPLETE TURBINE STAGE WITH 
UNIFORM AND NONUNIFORM INLET ~ 
TEMPERATURE PROFILES 
L2 
Ll 
LO 
... 
tt .9 I;:) 
o· .8 j ~ 
a:: ~ .7 
g g .6 1 
0: .4 
u 
ANNULAR CASCADE AT 
50% AXIAL CHORD 
PRESSURE 
SURFACE 
~ 
§ .5 t 
.... ~ . 
• 3 (. 
2 I I I I I 1 I I I 
• 11 18 19 20 21 22 23 24 25 
CIRCUHfERENTlAl POSITI~, 8, deq 
TtEORY 
--TSOOICI 
MERIDl 
-.- DODGE 
..... DENTON 
a LASER 
• STATIC 
TAPS 
WARM TURBINE 
LASER SURVEY 
WINDOW 
·BENCHMARK DATA FOR EVALUATION OF 
ADVANCED 3-D FLOW CODES FOR AERO~ 
. . 
DYNAMICS AND HEAT TRANSFER 
~ 
·w 
o 
(:-82-0868 
• SINGLE STAGE 20" TIP DIA 
• TIT 2000 TO 40000F 
• TIP 600 psia 
• WHEEL SPEED LIMIT 23,000 rpm 
• AIR FLOW 150 Ib/sec 
• GASPATH & COOLING-AIR TEMP. & PRESS. "."~~'~;"'\~].). ·'.('~i."'·' ~.; .. ·;:"9: ~~. ~~ 
• METAL TEMPERATURES AND HEAT FLUXES·:·'~~:g~:~:~.::?;·:.:::· .... , 
r-----. ______ .... ---i .. . :.~>.> .. ' " '" 
~~.: 
?~!,~ 
.''':'. 
.... 
" 
':', (:" 
131 
I-' 
eN 
N 
,iI •• 
National Aeronautics and . 
Space Administration TURBINE ENGINE HOT SECTION TECHNOLOGY 
Lewis Research Center 
PLASMA SPRAYED 
ENVI RONMENT 
COATING 
BUILT-IN GARDON TYPE 
HEAT FLUX SENSOR 
0.060 in diam 
LAMINATED TYPE 
HEAT FLUX GAGE 
0.250 in diam 
AND AIRFOIL 
CURVATURE 
~ III LASER OR 
NI\S/\ 
ELECTRON BEAM 
WELD 
ELECTRON BEAM 
OR DIFFUSION 
BOND JOINING 
LINE 
PLASMA SPRAYED 
CERAMIC MATERIAL 
ELECTRON BEAM 
. OR DIFFUSION BOND 
JOINING LINE ------' 
IN-HOUSE METHODS FOR THE FABRICATION 
OF AIRFOILS WITH HEAT FLUX SENSORS 
J-I 
W 
W 
• if •• 
Nallonal Aeronautics and 
Space Administration TURBINE ENGINE HOT SECTION TECHNOLOGY NI\SI\ 
Lewis Research Center 
ASPIRATED TEMP. 
PROBE, GAS 
ANALYSIS, OR 
PULSED THERMO-
COUPLE? 0/ //'- ~~~~I~ER 
r- TG 4 ASPIRATED 
~/ TARGET 
TG 
\ ' \ PROBE, GAS ANAI- TG 48 -. 
, YSIS, OR PULSED CERAMIC ' • , PI 
'~ERMOCOUP'-:7 PYROMETER '/1-. \ a TARGET7--" 
, ~ 
, 
" HEAT FLUX ~ 
GAGE----\ 
HEAT FLUX GAGES -.,. 
/ \ 
m 
/ , 
/ \ 
/ \ -
,- CERAMIC ~'-...~ 
¢ TARGET WITH" \,It(A ~ 
THERMOCOUPLE ,/ ~/ 
TG, 4.8, S LTG, 5 ASPIRATED 
TEMP. PROBE ~ , 
, 
TG -J \ 
\ o /~-- HOT GAS 
L TG ASPIRATED EXHAUST 
, 5.8, P 
, 4 
TEMP. PROBE 
LOCAL HOT GAS TEMPERATURE AND HEAT FLUX INSTRUMENTATION 
CD-81-IH'IO 
, 5.8, S 
'J-' 
W 
./:>0 
,II " 
TYPI CALAIRFOI L AEROTHERMODYNAM1C ANALY51 S METHOD 
HOT GAS 
. BOUNDARY 
CONDITIONS 
3-D BOUNDARY LAYER CODES 
STAN5~ FINITE DIFFERENCE 
BLAYER~ INTEGRAL 
. NODE 
GENERATOR 
INVISCID 
TSONIC~ 2D~ BLADE TO BLADE r-TW==O---:::OR=-=TH=RE=E:-:-D~I=ME:7N:-::S I=ON~A-:-L,,-~~ER IDL" 2D" HUB TO SHROUD 
~ NON-VISCOUS OR VISCOUS MIT" 3~" (THOMPKINS) 
FLOW CODE DENTON" 3D .. 
VISCOUS 
... 
-;. 
NANCY" 3D~ (DODGE) 
BOUNDARY LAYER HEA~ ~~PEPSI" 3D" (McDONALD) 
TRANSFER CODE . M.I. NT.". 3D" (McDONALD) 
P. D. THOMAS BEAM-WARMING CODE,,(UTSI 
_----.11-------.' WALL 
~ THREE-DIMENSIONAL TEMP. STRUCTURAl LIFE STEADY STATE AND AND ,ANALYSIS PREDICTION 
TRANSIENT CONDUCTION CODE PRESS. CODE CODE 
COOLANT-SIDE 
HEAT TRANSFER ~ 
'CORRELATIONS 
--INDUSTRY COOLANT-
SIDE CORRELATIONS 
--OPEN LITEl~TURE 
CORRELATIONS 
ONE-" TWO-" OR THREE-
DIMENSIONAL VISCOUS 
FLOW CaDE OR NEHIORK 
COilE. 
-- I NDUSTRV FLOW CODES 
--TACT}' IMPINGEMENT lNSERT 
--FCFC, FULL COVERAGE FILM COOLED 
-:.;TACJ2,,_ MULTI -PASS 
,..... 
W 
(J'I 
::"'~':1IOd TURBINE ENGINE HOT SECTION TECHNOLOGY I\JJ\S/\ 
~~c..-
HOST IN-HOUSE HIGH TEMPERATURE TURBINE TESTING 
ELEMENT 
CONDUCT HEAT TRANSFER 
STUDIES OVER THE RANGE 
OF REAL-EI~GINE. CONIlJTlONS 
WI TH NON-FI UI COOLED 
AIRFOILS 
VERIFICATION TEST OF 
THE NASA BASELINE 
RESEARCH TURBINE WITH 
AIRFOIL METAL TEMPER-
ATURE PREDICTION CODES 
MAP THE THREE-DIMENSIONAL 
FLOW FIELD IN A ROTATING 
PASSAGE WITH NO FILM 
COOLING 
MEASURE BI-AXIAL STRAIN 
DELTAS AT CRITICAL 
LOCATIONS ON STATIC AND 
ROTATING AIRFOILS AND 
THE DISK 
VERIFICATION TEST OF 
ADVANCED TECHNOLOGY 
COOLED TURBINE WITH 
AIRFOIL METAL TEMPER-
ATURE PREDICTION CODES 
FYI 82 
c=J 
WARH 
TURBINE 
c:=:I 
HG & HC 
c=J 
c:J 
HPT 
c:::J 
EXPECTED RESULTS 
~-REAL ENGINE VALUES OF LOCAL 
HOT GAS TEMPlRATURES, LOCAL 
HOT GAS-SIDE. HEAT TRANSFER 
COEFFICIENTS, AND LOCAL 
COOLANT-SIDE HEAT TRANSFER 
COEFFICIENTS FOR COMPARISON 
WITH ANALYTltAL MODELS AND 
CODES 
--ASSESSMENT OF AIRFOIL METAL 
TEMPERATURE PREDICTION CODES 
WITH THE NASA BASELINE RE-
SEARCH TUIlBINE 
--VERIFY THE INVISCID FLOW 
REGION OF THE PASSAGE, WHICH 
WILL HAVE BEEN ESTABLISHED IN 
WARM TURDINE TESTS AND 
ESTABLISH THE ACTUAL.VELOCI-
TIES IN THE SECONDARY FLOW 
REGIONS OF THE PASSAGE FOR 
COMPARISON WITH FLOW CODES 
--ASSESSMENT OF STRUCTURAL CODES 
WITH KNOWN THERMAL AND MECHANI-
CAL BOUNDARY tONDITIONS 
~ --ASSESSMENT OF AIRFOIL METAL 
TEMPERATURE PREDICTION CODES 
WITH AN ADVANtED TECHNOLOGY 
COOLED TURBINE 


Turbine heat transfer

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server