Modern jet engine design imposes extremely high loadings and temperatures on hot section components. A series of interdisciplinary modeling and analysis techniques which were specialized to address three specific components (combustor burner linings, hollow air-cooled turbine blades, and air-cooled turbine vanes) were developed and verified. These techniques will incorporate data as well as theoretical methods from many diverse areas, including cycle and performance analysis, heat transfer analysis, linear and nonlinear stress analysis, and mission analysis. Building on the proven techniques already available in these fields, the new methods developed will be integrated to predict temperature, deformation, stress, and strain histories throughout a complete flight mission

Roberts, M. L.

English

A series of interdisciplinary modeling and analysis techniques that were specialized to address three specific hot section components are presented. These techniques will incorporate data as well as theoretical methods from many diverse areas including cycle and performance analysis, heat transfer analysis, linear and nonlinear stress analysis, and mission analysis. Building on the proven techniques already available in these fields, the new methods developed will be integrated into computer codes to provide an accurate, and unified approach to analyzing combustor burner liners, hollow air cooled turbine blades, and air cooled turbine vanes. For these components, the methods developed will predict temperature, deformation, stress and strain histories throughout a complete flight mission

Mcknight, R. L.

NASA Technical Reports Server

Component-specific modeling

N87-11207
COMPONENT SPECIFIC MODELING
M. L. Roberts
General Electric Company
Aircraft Engine Business Group
I. INTRODUCTION
Modern jet engine design imposes extremely high loadings and temperatures on
hot section components. Fuel costs dictate that minimum weight components be
used wherever possible. In order to satisfy these two criteria, designers are
turning toward improved materials and innovative designs. Along with these
approaches, however, they must also have more accurate, more economical, and more
comprehensive analytical methods.
Numerous analytical methods are available which can, in principle, handle any
problem which might arise. However, the time and expense required to produce
acceptable solutions is often excessive. This program addresses this problem by
developing specialized software packages, which will provide the necessary
answers in an efficient, user-oriented, streamlined fashion. Separate
component-specific models will be created for burner liners, turbine blades, and
turbine vanes using fundamental data from many technical areas.
2. OBJECTIVE
The overall objective of this program is to develop and verify a series of
interdisciplinary modeling and analysis techniques which have been specialized to
address three specific hot section components: combustor burner liners, hollow
air-cooled turbine blades, and air-cooled turbine vanes. These techniques will
incorporate data as well as theoretical methods from many diverse areas,
including cycle and performance analysis, heat transfer analysis, linear and
nonlinear stress analysis, and mission analysis. Building on the proven
techniques already available in these fields, the new methods developed through
this contract will be integrated to predict temperature, deformation, stress, and
strain histories throughout a complete flight mission.
3. APPROACH
The work breakdown structure and tasks were discussed in detail at the Second
HOST Workshop last year, and will not be repeated here. Three major development
activities make up the Base Program. These are:
I. The Thermodynamic Engine Model,
2. The Thermomechanical Loads Model, and
243
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3. The Three Component Specific Models:
- Combustor liner,
- Turbine blade, and
- Turbine vane.
Thermodynamic Engine Model
The Thermodynamic Engine Model provides a decomposition/synthesis approach to
compute, at any point in a mission, the engine rotor speeds and the gas path
dynamic variables (velocity, temperature, pressure, density and Mach number) at
any station along the combustor liner and high pressure turbine flow path. With
this capability it is possible to synthesize the gas path variable history at any
station of the burner liner, blade or vane for a complete mission, without
computing the complete engine cycle for all mission points.
To develop the model, the engine cycle deck was run at 148 cycle points
covering the complete engine operating range shown in Figure I. From this data
an engine performance cycle map and an interpolation scheme were developed to
compute the gas path parameters at a chosen engine station, given the engine
operating conditions specified in terms of altitude, free stream Mach number and
engine power level. This also was discussed in more detail at last year's
Workshop. Results for a typical commercial airline flight: altitude, Mach
number, thrust, core rotor speed, compressor discharge temperature and turbine
inlet temperature, are shown in Figures 2 through 7. It is not necessary to run
complete missions. Segments can be run separately and joined to synthesize
alternate missions.
Thermomechanical Loads Model
The Thermomechanical Loads Model accepts as input variables the rotor speeds
and the gas path temperatures and pressures at one or more engine stations and
computes component metal temperatures and surface pressures. Again, a
decomposition/synthesis approach is used.
A typical commercial engine combustor liner, for which the loads model was
developed, is shown in Figure 8. The correlation for the burner liner
temperature, Tline r, was developed in terms of the cooling effectiveness
factor, nc: T -
4 Tliner
nc T4 - T3
244
where
T 3 = compressor discharge temperature, and
T4 = combustor exit temperature.
Using both engine test data and analytic results, a correlation of cooling
effectiveness with combustor length was developed for a single combustor panel
(Figure 9). For a realistic picture of liner temperatures, both average
temperature and the hot streak temperature associated with the fuel nozzles were
correlated. The effects of pressure and of altitude on cooling effectiveness
were examined separately. Both were small enough to be neglected. To account
for the temperature gradient through the metal thickness, an expression was
derived from cooling effectiveness, compressor discharge temperature and
pressure, and combustor exit temperature. By correlating test data against
combustor pressure, a representative expression for temperature gradient through
the metal thickness was obtained at five typical points along the length of the
combustor liner. This completed the work on the combustor liner Thermomechanical
Loads Model.
A common approach was taken for the cooled high pressure turbine blade and
vane. Cooling effectiveness was again used as the correlating factor. The
initial step was to correlate cooling effectiveness as a function of engine
operating conditions at the 50% span station. Cross sections at 50% span of the
typical blade and vane configurations are shown in Figures I0 and II. A typical
correlation of cooling effectiveness with gas path temperature and power level at
the 50% span station on the turbine vane is shown in Figure 12. Both test data
and analytical predictions are shown. The trend of the data is quite similar to
the combustor data; however, there is some disagreement between the analytical
and test values of nc. This was not unexpected because of: (I) differences
between nominal values of design variables (Tqas, Tcoolant, wall thicknesses
and thermal properties) and actual values; (27 uncertainties in calculating the
gas side and coolant side heat transfer coefficients, the radiation heat flux,
film cooling effectiveness, etc., (3) measured temperature errors due to
uncertainties in thermocouple measurements, flow checking measurements, etc. As
a result the test data was used to validate the trend lines, and the analytic
predictions were used for the correlation.
From data of the type shown in Figure 12, a correlating equation was
developed to predict cooling effectiveness at the 50% span station as a function
of operating conditions:
l-n c T3 0.04 T 0.04
_ ) ( 4.1 ,Ref )
l-nc,Re f (T3,Re f T4o I
245
where nc,Ref, T3 Ref and T4.1 .Ref denote the reference values and
conditions of Figures lO and II.
To predict the metal temperatures at other points on the blades, the cooling
effectiveness factors at other spanwise stations were correlated against the
values at the 50% span station, 50%- This was done at eight chordwise
stations on both the suction and pressure surfaces. Thus, using the gas path
temperatures T3 and T4. l for any engine operating condition, the reference
data shown in Figures lO and II, and the correlating equation for nc, the
Thermomechanical Loads Model can be used to predict the metal temperatures over
the complete blade and vane surfaces.
To establish the general procedure for predicting the static gas pressure
distributions along the airfoil surfaces of the blade and vane, typical design
gas pressure distributions were collected and normalized. Figure 13 shows the
typical turbine normalized vane gas static pressure distribution.
With the completion of the Thermomechanical Loads Model, the capability
exists for synthesizing a mission, determining the hot section flow path gas
properties for the complete mission, and, from these, the metal temperature and
pressure histories for the combustor and turbine blade and vane.
Component Specific Model Development - Geometric Modeling
The approach taken to model the three components is to select typical
components, identify the key input parameters and develop master regions from
these parameters. The finite element mesh can then be overlaid on each master
region. Figure 14 shows a typical combustor nugget, some of the physical input
parameters, and the master region definition based on those parameters. Figure
15 shows representative 2D and 3D models generated from these master regions.
Modeling of the other two components, the blade and vane, is currently
progressing.
4. CONCLUSION
When completed, this program will provide a non-linear stress analysis system
for hot section engine parts that will allow the component designer to evaluate
quickly the effects of mission variations on component life. It will be easy to
use, cost effective, and make a significant contribution to assessing hot section
durability.
246
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247
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FIGURE4, TH__RUSTVS, TIME
248
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FIGURE5, CORE ROTOR SPEED VS, TIMF
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FIGURE 6, COMPRESSORDISCHARGETEMPERATUREVS, TIME
249
TURBINE
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FIGURE 7. TURBINE INLET TEMPERATURE VS, TIME
i 2 5 6
\\ I 1. DomeUpstream
"_ t 2. Dome Downstream
3 3. Inner Passage, Upstream
4. Inner Passage, Downstream
5. Outer Passage, Downstream
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FIGURE 8, CF6-50 ROLLED RING COMBUSTOR
250
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FIGURE9. COOLINGEFFECTIVENESSDISTRIBUTION,PANEL 7
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251
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253
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254


Component specific modeling

A series of interdisciplinary modeling and analysis techniques that were specialized to address three specific hot section components are presented. These techniques will incorporate data as well as theoretical methods from many diverse areas including cycle and performance analysis, heat transfer analysis, linear and nonlinear stress analysis, and mission analysis. Building on the proven techniques already available in these fields, the new methods developed will be integrated into computer codes to provide an accurate, efficient and unified approach to analyzing combustor burner liners, hollow air-cooled turbine blades and air-cooled turbine vanes. For these components, the methods developed will predict temperature, deformation, stress and strain histories throughout a complete flight mission

The overall objective of this program is to develop and verify a series of interdisciplinary modeling and analysis techniques which have been specialized to address three specific hot section components. These techniques will incorporate data as well as theoretical methods from many diverse areas, including cycle and performance analysis, heat transfer analysis, linear and nonlinear stress analysis, and mission analysis. Building on the proven techniques already available in these fields, the new methods developed through this contract will be integrated to provide an accurate, efficient, and unified approach to analyzing combustor burner liners, hollow air cooled turbine blades, and air cooled turbine vanes. For these components, the methods developed will predict temperature, deformation, stress, and strain histories throughout a complete flight mission

COMPONENT-SPECIFIC MODELING* 
M. L. Roberts 
General Electric Company 
Introduction 
Modern jet engine design imposes extremely high loadings and temperatures on 
hot section components. Fuel costs dictate that minimum weight components be used 
wherever possible. In order to satisfy these two criteria, designers are turning 
toward improved materials and innovative designs. Along with these approaches, 
however, they must also have more accurate, more economical, and more comprehensive 
analytical methods. 
Numerous analytical methods are available which can, in principle, handle any 
problem which might arise. However, the time and expense required to produce 
acceptable solutions is often excessive. This program addresses this problem by 
setting out a plan to create specialized software packages, which will provide the 
necessary answers in an efficient, user-oriented, streamlined fashion. Separate 
component-specif ic models wi 11 be created for burner 1 iners, turbine blades, and 
turbine vanes using fundamental data from many technical areas. The methods 
developed will be simple to execute, but they will not be simple in concept. The 
problem is extremely complex and only by a thorough understanding of the details 
can the important technical approaches be extracted. The packaging of these 
interdisciplinary approaches into a total system must then conform to the modular 
requirements for useful computer programs. 
Objective 
The overall objective of this program is to develop and verify a series of 
interdisciplinary modeling and analysis techniques which have been specialized to 
address three specific hot section components. These techniques will incorporate 
data as well as theoretical methods from many diverse areas, including cycle and 
performance analysis, heat transfer analysis, 1 inear and nonl inear stress analysis, 
and mission analysis. Building on the proven techniques already available in these 
fields, the new methods developed through this contract will be integrated to 
provide an accurate, efficient, and unified approach to analyzing combustor burner 
liners, hollow air-cooled turbine blades, and air-cooled turbine vanes. For these 
components, the methods developed will predict temperature, deformation, stress, 
and strain histories throughout a complete flight mission, 
Background 
This program, to a great extent, will draw on prior experience. This base of 
experience is invaluable for understanding the highly complex interactions among 
all the different technical disciplines as well as for estimating the importance of 
different engine parameters. In particular, there are four specific areas in which 
experience will be especially beneficial. 
*This work was done under NASA Contract NAS3-23687. 
https://ntrs.nasa.gov/search.jsp?R=19860002048 2020-03-20T16:13:12+00:00Z
First, with the recent increases in fuel costs, greater emphasis has been 
placed on more accurate solutions for stresses and strains in order to understand 
and improve the durability and life of hot section components. Conventional linear 
elastic analyses are no longer sufficient; instead, they now provide the boundary 
values for more refined creep and plasticity calculations. These nonlinear 
analyses are now performed routinely as part of the design process at Genera1 
Electric. This extensive experience with these plasticity and creep methods will 
contribute directly to developing component specific models. 
Second, advances in 3-D modeling capabi 1 ity are being achieved by the concepts 
developed under the NASA-supported ESMOSS program. ESMOSS concepts wi 11 provide 
the basis to develop an efficient modeling system for geometric and discretized 
models of engine components. 
Third, the NASA-funded Burner Liner Thermal/Structural Load Model ing Program 
will contribute strong support to this program. The specific area addressed, 
transfer of data from a 3-D heat transfer analysis model to a 3-D stress analysis 
model, will provide the background and framework for the data interpolation 
required for all thermomechanical models in this contract. 
Fourth, over the past 10 years, General Electric has developed internally a 
family of computer programs: LASTS, OPSEV, and HOTSAM. These programs all have the 
common thread of using selected points from cycle data, heat transfer, and stress 
analyses, and a decomposition/synthesis approach to produce accurate values of 
temperature, stress, and strain throughout a mission. These programs are totally 
consistent with the overall objectives of this program, and represent a proven 
technology base upon which the component specific models will be developed. 
Significant advances to be made are the inclusion of nonlinear effects and the 
introduction of improved modeling and data transfer techniques. 
Approach 
The program is organized into nine tasks which can logically be separated into 
two broadly parallel activities (Figure 1). On the right of Figure 1 we have the 
Component Specific Thermo-Mechanical Load Mission Modeling path. Along this path a 
Decomposition/Synthesis approach wi 11 be taken. In broad terms this means 
developing methods to generate approximate numerical models for the engine cycle 
and the aerodynamic and heat transfer analyses needed to provide the input 
conditions for hot parts stress and life analysis. 
The left path, Component Specific Structural Modeling will provide the tools to 
develop and analyse finite element nonlinear stress analysis models of combustor 
liners and turbine blades and vanes. These two paths are shown in more detail in 
Figures 2 and 3 .  
Software Development, Task IV consists of planning and writing the computer 
programs for both paths, with the necessary interconnections, using a structured, 
top down approach. 
In the Thermomechanical Load Mission Modeling portion of the program (Figure 2) 
we will develop, in Task 111, a Thermodynamic Engine Model which will generate the 
engine internal flow variables for any point on the operating mission. The method 
for doing this is described below, Task V will develop techniques to decompose 
flight missions into characteristic mission segments. In Task VII a Thermo- 
mechanical Mission Model will be developed. This will use the flow variables from 
the Thermodynamic Model to determine metal temperature and pressure distributions 
for a representative combustor liner and turbine blade and vane. 
Individual tasks for the Structural Modeling activity are shown in Figure 3. 
The requirements of Software Design, Task 11, w ~ l l  be factored into Task VI, the 
evaluation of the structural analysis methods which were selected for evaluation in 
Task I. Task VIII will provide the capability for structurally modeling current 
state-of -the-art combustor 1 iners and hol low turbine blades and vanes, given the 
defining dimensional parameters. These parameters will be chosen to facilitate 
parametric studies. 
The component s ecif ic models will be developed in two steps. In the first a e geometric model wil be defined. In the application of the Component Specific 
Modeling Pro ram this data will then be transferred to the Thermomechanical Load 
Mission Mode 9 to provide the geometry for determining component pressures and 
temperatures. Thus, a data transfer link will be developed to do this in Task IV, 
Software Development. The capability for generating from the geometric model a 
discretized, finite element model will also be part of Task VIII. At this point 
another link between the two paths will be needed to transfer the component 
temperatures and pressures from the Thermomechanical Load model to the finite 
elevent model, interpolating the data as needed to define nodal temperatures and 
pressures. This also will be completed in Task IV. 
The final function in Task VIII will be the development of component-specific 
stress analysis models for the three components to perform cyclic elastic, plastic 
and creep analyses using loading conditions defined by the Thermomechanical Load 
Models Progress. 
At this time considerable progress has been made on the Thermodynamic Engine 
Model. The model is being developed as a simple calculational tool which will take 
as inputs the three variables, altitude (h), Mach number (M) and power level (PL) 
for the allowed flight map of an engine, as shown in Figure 4. In addition, 
ambient temperature deviations from the standard atmosphere, airframe bleed air 
requirements and engine deterioration can also be included as part of the input to 
the Thermodynamic Model. For each input condition, specified by h, M and PL the 
Thermodynamic Engine Model wi 11 calculate gas weight flow ( i )  , temperature (t) and 
pressure (p) at selected aerodynamic engine stations, as needed to determine 
component thermal loadings. These stations are shown in Figure 5. 
The technique for develo ing a Thermodynamic Engine Model is shown in Figures 6 7 and 7. The engine to be ana yzed must be defined thermodynamically by an engine 
cycle deck (computer program) which can be run to generate the internal flow 
variables at the chosen aerodynamic stations (Figure 6). To encompass the complete 
en ine operating map (Fi ure 4), 148 operating points are chosen and i ,  t and p are 
ca 9 culated using the cyc 9 e deck for the selected stations, as well as N1 and N 2 ,  
the fan and core speeds. From this station data an Engine Performance Cycle Map is 
constructed. This is essentially a set of three-dimensional data arrays which map 
the station data ( i ,  t, p, N 1  and N2) on to the engine operating map (Figure 4). 
Given an arbitrary operating point defined by h, M and PL it is then, in principle, 
possible to interpolate on the Engine Performance Cycle Map to determine station 
data. In practice the station parameters are nonlinear functions of the input 
parameters and considerable effort was needed to develop these mult i-dimensional 
interpolations, The computer program used to generate the Engine Performance Cycle 
Map from the engine cycle desk output has been developed as part of Task 111. The 
functioning of the Thermodynamic Engine Model is shown in Figure 7. Given an 
engine mission, as shown schematically in Figure 8 it can be defined by values of 
the input variables h, M and PL at selected times through the mission. Using these 
input variables and the Engine Performance Cycle Map an Interpolation Program, now 
being developed in Task I11 of this program, will calculate engine station 
parameters throughout the mission (Figure 7). These are then used to define 
Station Mission Profiles of G, t, p, N1, and N2, as functions of time at each 
aerodynamic station. These Station Mission Profiles are become the input to the 
Thermomechanical Engine Model. 
The Thermomechanical Model is less well developed at this time than the 
Thermodynamic Model. Its form will be based on types of correlations previously 
developed within General Electric. Figure 9 shows a representative correlation for 
a turbine vane. Metal temperatures at various points on the vane TVA are 
correlated in terms of a vane overall cooling effectiveness, TI, and station gas 
temperatures T3 at compressor discharge, and T4 at combustor discharge. Using the 
Station Mission Profiles it will be possible to calculate the temperatures at 
selected locations on each component as functions of time, given the input 
parameters h, M and PL that define the engine mission. These then will provide the 
boundary conditions for the component stress analyses. 
On the Component Specific Structural Modeling path, concepts have been defined 
and are being implemented. Additional evaluations are needed, however, before they 
can be presented for discussion. 
Conclusion 
When completed this program will provide a stress analysis system for hot 
section parts that will allow the component designers to evaluate quickly the 
effects of mission variations, be easy to use, cost effective, and make a 
significant contribution to assessing hot section durability. 
FIGURE 1.  COMPONENT S P E C I F I C  MODELLING - BASE PROGRAM 
MISSION HIDEL 
------- 
DECOMWS IT ION/SYNTHESIS 
FIGURE 2 .  COMPONENT S P E C I F I C  
THERMOMECHANLCAL LOAD M I S S I O N  MODELLING 
IAxu 
DESIGN STRUCTURAL ANALYSIS 
SOFTWARE ARCHITECTURE 
-
CTURAL ANALYSIS PONENT SPEC I F 1  C MODELING 
HODS EVALUATION -----SELECT COMPONENTS FOR 1 I MODEL DEVELOPMENT I 
DEVELOP GEOMETRIC I H O D E L I N G L D I S P L A Y  I 
DEVELOP STRUCTURAL 
ANALYSIS CAPABILITY 
FIGURE 3. COMPONENT SPECIF IC  STRUCTURAL MODELING 
FIGURE 4. ENGINE OPERATLNG MAP 
Temperature 5 
0 AMBIENT 
2 ENGINE INUT 
2.2 FAN UP (BYPASS) s r a m m  INLET 
IC  FAN HUB (LP COWRESSOR) SREAM I N U T  
2.3 FAN DISCHARGE 
2.4 LP COMPRESSOR DISCHARGE 
2D LP COMPRESSOR DISHCARGE ILEED PORT EXIT 
2.5 BYPASS SREAM MIXING PLANE 
2.6 IYPASS DUCT INLET (AFTER MIXING) 
2.8 IYPASS D u c r  JET NOZZLE THROAT 
2..9 BYPASS W C T  X T  NOZZLE EXIT 
(COMLETE EXPANSION) 
HP COMPRESSOR INLET 
W COMPRLSSOR D l X W G E  
HP TURBINE IST-STAGE NOZZLE INLET (W/O COOLING FLOW) 
HP TURBINE ROTOR INLET (W/ COOLING FLOW) 
HP TUIINE M K W G E  (W/O COOLING FLOW) 
HP TURUNE DISC)(ARGE IW/ COOLING FLOW) 
u TURBINE INLET (w/ COOLING FLOW) . 
LP TURBINE DISCHARGE (W/O COOLING FLOW) 
LP R I R I I N E  M X W G E  iW/COOLINO FLOW 
TE'EMPERATUE FAN - - - - - PRESSUKE FAN - - - - - - 
CORE 
F%RE 5. 6 AERODYNAMIC STATIONS 
TR42 l -1  
1 4 8  
ENGINE ENGINE 
OPERTG, PERF0 W N C E  
POINTS CYCLE MAP 
1, HJ To. 1 4 8  CASES 
PWR, LVLJ L. TJ PJ HI, N;, 
CUST, BLDJ AT EACH 
DETRN, LVL. STAT ION 
FIGURE 6. THERMODYNAMIC ENGINE MODEL 
CYCLE MAP GENERATION 
FIGURE 7, THERMODYNAMIC ENGINE MODEL 
Time - Minutes 
FIGURE 8. TYPICAL FLIGHT CYCLE 
FIGURE 9 .  TURBINE VANE COOLING EFFECTIVENESS 


Accomplishments are described for the second year effort of a 3-year program to develop methodology for component specific modeling of aircraft engine hot section components (turbine blades, turbine vanes, and burner liners). These accomplishments include: (1) engine thermodynamic and mission models; (2) geometry model generators; (3) remeshing; (4) specialty 3-D inelastic stuctural analysis; (5) computationally efficient solvers, (6) adaptive solution strategies; (7) engine performance parameters/component response variables decomposition and synthesis; (8) integrated software architecture and development, and (9) validation cases for software developed

Component specific modeling

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NASA Technical Reports Server

NASA Technical Reports Server

NASA Technical Reports Server

NASA Technical Reports Server

NASA Technical Reports Server