Three dimensional nonlinear finite element heat transfer and structural analyses were performed for the first stage high pressure fuel turbopump (HPFTP) blade of the space shuttle main engine (SSME). Directionally solidified (DS) MAR-M 246 and single crystal (SC) PWA-1480 material properties were used for the analyses. Analytical conditions were based on a typical test stand engine cycle. Blade temperature and stress strain histories were calculated by using the MARC finite element computer code. The structural response of an SSME turbine blade was assessed and a greater understanding of blade damage mechanisms, convective cooling effects, and thermal mechanical effects was gained

Abdul-Aziz, Ali

Thompson, Robert L.

English

NASA Technical Reports Server

N88-22404
THERMAL-STRUCTURAL ANALYSES OF SPACE SHUTTLE
MAIN ENGINE (SSME) HOT SECTION COMPONENTS
Ali AbduI-Aziz* and Robert L. Thompson
Structural Mechanics Branch
NASA Lewis Research Center
ABSTRACT
Three-dimensional nonlinear finite-element heat transfer and structural analy-
ses were performed for the first-stage high-pressure fuel turbopump (HPFTP)
blade of the space shuttle main engine (SSME). Directionally solidified (DS)
MAR-M 246 and single crystal (SC) PWA-1480 material properties were used for
the analyses. Analytical conditions were based on a typical test stand engine
cycle. Blade temperature and stress-strain histories were calculated by using
the MARC finite-element computer code (MARC Analysis Research Corporation,
1980).
This study was undertaken to assess the structural response of an SSME turbine
blade and to gain greater understanding of blade damage mechanisms, convective
cooling effects, and thermal-mechanical effects (Kaufman, 1984). Other objec-
tives are also to address problems of calculating the thermal response of
high-temperature gas-path components, such as turbopump blades in space power
propulsion systems, and to evaluate the stress-strain response at the blade
critical location for life prediction purposes. Another purpose is to exercise
nonlinear finite-element computer codes on anisotropic SSME components for a
typical mission cycle.
Thermal environment on the airfoil was predicted by running a boundary layer
analysis using STAN5 code (Gaugler, 1981) to generate the film coefficients
needed for the heat transfer analysis. For the complete blade, the heat trans-
fer conditions around the platform and shank regions were provided by Lockheed
Missiles & Space Company, Huntsville Research and Engineering Center (1983).
Additional assumptions also had to be made to complete the boundary conditions.
Transient results were obtained by scaling steady-state heat transfer coeffici-
ents based on transient flow and temperature. More details regarding the heat
transfer analysis are given by Abdul-Aziz (1987).
Elastoplastic analyses have been conducted for the HPFTP blade with the MARC
code. Plastic strains calculations were based on incremental plasticity theory
using von Mises yield criterion, the normality rule, and a kinematic hardening
model. The material elastoplastic behavior was specified by the yield strength
and work hardening properties in the longitudinal direction; transverse proper-
ties were not available. Approximately 2 million words of core storage on the
*Sverdrup Technology, Inc., Lewis Research Center Group, NASA Lewis
Research Center.
2-255
https://ntrs.nasa.gov/search.jsp?R=19880013020 2020-03-20T06:32:34+00:00Z
Cray XMPcomputer were needed to run the problem. Analysis required about
4 hours of Cray XMPcentral processor unit (CPU) time. About half of that was
needed for the airfoil problem.
2-256
TURBINEBLADEFINITE-ELEMENTMODEL
In the past, first-stage turbine blades in the high-pressure fuel turbopump(HPFTP)of the space shuttle main engine have suffered cracking in the blade
shank region and at the leading edge of the airfoil immediately above the plat-
form. A finite-element model for a complete blade has been constructed with
eight-node, solid, isoparametric elements. Another model is shown for an
alternate airfoil design which is also undergoing thermal-structural analyses.
The complete-blade finite-element model consists of 1025 elements and 1575
nodes. The airfoil portion contains 360 elements and 576 nodes, while the
alternate airfoil design includes 546 elements and 840 nodes. The MARCcode
and PATRAN-Ggraphic package were used to construct these finite-element
models.
I
FINITE-ELEMENT MODEL
LEADING
EDGE-,,,
PLATFORM_'l AIRFOIL
ALTERNATE AIRFOIL
FIRST-STAGE BLADE
CD-88-31703
2-257
MISSION CYCLE FOR ANALYSIS
The mission cycle used for the analyses in terms of measured and corrected
turbine inlet gas temperature, gas pressure, and blade rotational speed is
shown below. This cycle is typical of a factory test of the engine; it is
also reasonably representative of a flight mission except for the fore-
shortened steady-state operating time. Since the major factor inducing
cracking is the transient thermal stresses caused by the sharp ignition and
shutoff transients, deletion of the steady-state portion is acceptable.
2000
TURBINE 1200
INLET GAS
TEMPERATURE,
°C 400
-400
O
TURBINE
INLET
PRESSURE,
MPa
5O
40
30
20
10
MEASURED
CORRECTED
0
STEADY STATE-,, TURBINE
INLET GAS
1 2 3 4 5
TIME, sec
1000
600
STARTUP TRANSIENT
200
-200
28 29 30 31 32
TIME, sec
CUTOFF TRANSIENT
40 000 --
BLADE
ROTATIONAL 20 000
SPEED,rpm
I I I
8 16 24 32 0
TIME, sec
I I I
8 16 24 32
TIME, sec
CD-88-31704
t
33
2-258
AIRFOIL TEMPERATURE DISTRIBUTION 
Metal temperatures obtained from MARC heat transfer analysis for the two 
startup transient spikes are shown below. 
highest temperature occurred at the trailing edge near the tip where the air- 
foil is the thinnest. For the second ignition spike, the maximum temperature 
was at both the leading and trailing edges. 
enced a uniform temperature distribution except near the base because of colder 
boundary conditions in the region. 
the leading and trailing edges and was due to high gas-path pressures, which 
resulted in high heat transfer coefficients in those areas. The temperature 
reached 1142 O C  during the first spike and 7 7 7  O C  at cruise. 
For the first ignition spike, the 
During cruise the airfoil experi- 
The maximum temperature location was at 
MAXIMUM 
TEMPERATURE7 
I 
AX. TEMP. = 1142 "C \w (2087 "F) 
MIN. TEMP. = 800 'C 
(1470 'F) 
FIRST SPIKE (0.6 sec) 
MAX. TEMP. = 702 "C 
(1296 'F) 
MIN. TEMP. = 472 'C 
(882 'F) 
SECOND SPIKE (1.3 sec) 
'URE 
MAX. TEMP. = 777 "C 
(1430 'F) 
MIN. TEMP. = 602 'C 
(1115 'F) 
LST '88 cQ-0&3:771 
2-259 
AIRFOIL TRANSIENT TEMPERATURE RESPONSE
The thermal response predicted from the finite-element analysis shown below
revealed that the leading and trailing-edge bases of the airfoil are the
hottest locations. This was expected because of the presence of high heat
transfer coefficients in the region. However, the corrected temperatures
(which are higher than those measured by the poor thermocouple response) gave
rise to 28 percent higher leading-edge temperatures and a i00 percent higher
temperature between the midspan surface and the leading-edge base. The coldest
spot was always at the airfoil base because of colder boundary conditions.
METAL
TEMPERATURE,
°C
1200 -
1000
800
600
400
200
0
-200_
1000 -
800
600
400 -
200
-20_2_
j
1 I I
1 2 3 4
CORRECTED
STARTUPTRANSIENT
I I
27 20 29 30 31 32
TIME, sec
CUTOFFTRANSIENT
TRAILING-EDGETIP
TRAILING-BASE
LEADING-EDGEBASE
I
L
I
0
j J,"/,"
i i i
1 2 3 4 5
TIME, sec
MEASURED
STARTUPTRANSIENT
CD-88-31705
2-260
AIRFOIL RADIAL STRESS DISTRIBUTION AT CRUISE 
The radial stress components (i.e, along the span) are shown for the elasto- 
plastic MARC airfoil analyses of directionally solidified (DS) MAR-M 246 and 
single crystal (SC) PWA-1480. Incremental loading included centrifugal and 
gas-pressure loads and metal temperature distributions as calculated from the 
heat transfer analysis. Maximum stresses reached 119 ksi for the SC material 
and 110 ksi for the DS material during the cruise portion of the mission cycle. 
The radial component of the strain ranges were 2.1 percent for the MAR-M 246 
and 1.57 percent for the PWA-1480. 
MAX. TENSILE STRESS = 76 ksi 
MIN. COMPRESSIVE STRESS = -44 ksi 
DS MAR-M 246 
MAX.TENSILE STRESS = 68 ksi 
MIN. COMPRESSIVE STRESS = -53 ksi 
1 SC PWA-1480 
SUCTION SIDE PRESSURE SIDE 
C 3 - 8 6 - 2 ' 7 E  
2-26 1 
COMPLETE-BLADE TEMPERATURE CONTOURS 
Temperature contours on the surface of the blade are shown for the first tem- 
perature spike and at cruise. As expected, the hottest spot of about 1110 "C 
occurred at the trailing-edge tip of the airfoil during the first temperature 
overshoot. 
junction of the airfoil, platform, and shank, where a steep temperature gradi- 
ent is clearly shown. 
mixture in the region. 
the platform and shank regions may have affected the local heat transfer 
predictions. 
Uniform temperature dominated most of the airfoil except at the 
This is because of the effects of the cold and hot gas 
A lack of reliable data on the thermal environment in 
MAX. TEMP. = 772 "C 
(1422 O F )  
MIN. TEMP. = 472 "C 
(882 O F )  
CRUISE 
MAX. TEMP. = 1179 "C 
(2150 O F )  
MIN. TEMP. = 589 'C 
(1092 "F) 
FIRST SPIKE 
3-36? 
COMPLETE-BLADE RADIAL STRESS DISTRIBUTION AT CRUISE 
The results of elastoplastic analyses with the MARC code are shown below for 
the radial components of stress on the complete-blade structure for both 
DS MAR-M 246 and SC PWA-1480. Considerable difference in structural response 
has been observed in analyzing the complete blade rather than the airfoil 
alone. This is mainly due to the uncertainties in the prescribed boundary 
conditions and assumptions made at the shank-platform interface. However, any 
difference in life for the two materials would come from their fatigue 
characteristics. While the fatigue resistances are not well defined, results 
indicate thermal low-cycle fatigue lives of several thousand cycles. 
c 
SUCTION SIDE 
LST '00 
MAX. TENSILE STRESS = 118 ksi 
MIN. COMPRESSIVE STRESS = -40 ksi 
DS MAR-M 246 
MAX.TENSILE STRESS = 110 
MIN. COMPRESSIVE STRESS 
ksi 
= -40 
SC PWA-1480 
_ _  _- 
PRESSURE SIDE 
-- .i 
CD-88-31708 
2 - 2 6 3  
REFERENCES
Abdul-Aziz, A., Tong, M., and Kaufman, A., 1987, "Thermal Finite-Element
Analysis of an SSMETurbine Blade," NASATM-100117.
Gaugler, R.E., 1981, "SomeModification to, and Operational Experiences With,
The Two-Dimensional, Finite-Difference, Boundary-Layer Code, STANS,"NASA
TM-81631.
Kaufman, A., 1984, "Development of a Simplified Procedure for Cyclic
Structural Analysis," NASATP-2243.
Lekhnitskii, S.G., and Brandstatter, J.J., eds., 1963, Theory of Elasticity of
an Anisotropic Elastic Body," Holden-Day, Inc., pp. 24-25.
Lockheed Missiles & Space Company, Huntsville Research and Engineering Center,
1983, "Space Shuttle Main Engine, Powerhead Structural Modeling, Stress
and Fatigue Life Analysis," NASA Contract NAS8-34978, NASA CR-170999.
MARC Analysis Research Corporation, 1980, MARC General Purpose Finite Element
Analysis Program, Vol. A: User Information Manual; Vol. B: MARC Element
Library; Vol. F: Theoretical Manual, Palo Alto, CA.
Milligan, W. Walter, Jr, 1986, "Yielding and Deformation Behavior of The
Single Crystal Nickel-Base Superalloy PWA 1480," NASA CR-175100.
2-264


Thermal-structural analyses of Space Shuttle Main Engine (SSME) hot section components

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