The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are described. The expert system incorporates a JANNAF standard reference computer code to determine rocket engine performance and a state-of-the-art finite element computer code to calculate the interactions between propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb heuristics were incorporated for the hydrogen-oxygen coaxial injector design, including a minimum gap size constraint on the total number of injector elements. One-dimensional equilibrium chemistry was employed in the energy release analysis of the combustion chamber and three-dimensional finite-difference analysis of the regenerative cooling channels was used to calculate the pressure drop along the channels and the coolant temperature as it exits the coolant circuit. Inputting values to describe the geometry and state properties of the entire system is done directly from the computer keyboard. Graphical display of all output results from the computer code analyses is facilitated by menu selection of up to five dependent variables per plot

Davidian, Kenneth J.

English

NASA Technical Reports Server

NASA Technical Memorandum 102373
A Rocket Engine Design
Expert System _?
(NACA-TH-102373) A R_CKET _NGINE DFSIGN
E×PERT SYSTEM (NASA. Lewis Research
C_nter) II p CSCL 21H
G5/20
......... _ ___ :_
NoO-IOI?2
Uncl as
0234o94
Kenneth J. Davidian
Lewis Research Center
Cleveland, Ohio
.--_S T .......
Prepared for the __
26th JANNAF Combustion Meeting
--_:--_ Pasadena,- C_-l_fo-rn]a, O_0_i_f 23-27, 1989
https://ntrs.nasa.gov/search.jsp?R=19900000856 2020-03-20T00:23:14+00:00Z

A Rocket Engine Design Expert System
Kenneth J. Davidian
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio 44135-3191
ABSTRACT
The overall structure and capabilities of an expert system designed to evaluate rocket en-
gine performance are described. The expert system incorporates a JANNAF standard refer-
ence computer code to determine rocket engine performance and a state-of-the-art finite ele-
ments computer code to calculate the interactions between propellant injection, energy re-
lease in the combustion chamber, and regenerative cooling heat transfer. Rule-of-thumb
heuristics have been incorporated for the hydrogen-oxygen coaxial injector design, including
a minimum gap size constraint on the total number of injector elements. One-dimensional
equilibrium chemistry was employed in the energy release analysis of the combustion cham-
ber. A three-dimensional conduction/one-dimensional advection analysis is used to predict
heat transfer and coolant channel wall temperature distributions, in addition to coolant tem-
perature and pressure drop. Inputting values to describe the geometry and state properties of
the entire system is done directly from the computer keyboard. Graphical display of all output
results from the computer c&le analyses is facilitated by menu selection of up to five depen-
dent variables per plot.
INTRODUCTION
The process of designing a rocket engine involves many complex domains of technical ex-
pertise (figure 1). Each discipline has a unique set of computer codes, a unique set of past
experimental experience, and a specific set of constraints which guide the component design
process. Overall engine size and performance parameters are determined by an analyst who
is concerned with the overall mission and vehicle. The injector is often selected using exist-
ing experience based on past performance. Regenerative cooling channels and the thrust
chamber design are also chosen for good historical performance reasons. Other areas of en-
gine design include performance analysis of the expanding gases, design of a high performing
expansion nozzle, predicting the atomization process of the chosen injector, and evaluating
the combustion and injection processes for the possibility of combustion instability.
Designing a rocket engine also requires the solution of many multidisciplinary issues. Ac-
curate design of rocket engine components assumes that an open channel of communication
exists between the resident experts in each of the knowledge domains, and relies upon an it-
erative process to arrive at a solution. The injector designer needs to know the engine pa-
rameters and the propellant injection temperatures to design the elements. To do his job, the
chamber designer requires the propellant injection temperatures and chamber geometry data
to calculate the heat transfer to the regenerative cooling channels. The amount of energy in-
jected into the chamber, transferred through the chamber wall to the coolant, and then reintro-
duced into the chamber is determined by successively using improved estimates as calculat-
ed by the computer codes.
The computer codes in use by the designer of each component have always been complex
to run and to understand due to the necessity of modeling the physical phenomena as closely
as possible. The task of running single-point designs, which are commonly derived by slight-
ly varying previouslyrun cases,is far easierthandevelopinginputdatasetsfor a designfrom
scratch.Evenafter thecomputerprogramacceptsthe inputdata,theusermustbecarefulin
interpretingtheresults.
As aresultof thevariousnecessaryareasof expertise,therequiredinteractionbetween
theseareas,andthecomplexityof eacharea'scomputationaltools,someless-than-rigorous
approachesto rocketenginedesignhavebeenestablishedin eachdiscipline.Hard-earned
experienceshavebeencondensedto simplified "rules-of-thumb," thelackof openchannels
of communicationbetweendisciplineshasresultedin "open-loop" approachesto enginede-
sign,andcomputercodeswhichmakea compromisebetweentechnicalthoroughnessand
ease-of-useandunderstandingarerelieduponto a greatextent.
To providearigoroussolutionto theproblemof rocketenginedesign,theobstaclesmen-
tionedaboveneedto beeliminated.All disciplinesmustberepresented,if not in acomplete-
ly thoroughfashion,thenby their "rules-of-thumb" whichcanberelieduponto provideengi-
neeringapproximationsto specificquestions.Complexcodesmustgive theengineermore
thanhard-to-readtablesof output.Theymusthaveauser-interfacewhichmakesthecodes
easyto use,allowsfor thesharingof input databetweenmorethanonecode,andprovides
theusereasy-to-understandoutput information.Thecomputercodesneedto interactwith
eachother,allowing for iterationtowmdsasolutionanddesignoptimization.Finally, the
codesneedto becollectedin aframeworkwhichallows for futuregrowthof capability,up-
gradesof currentfunctions,andinclusionof newcomputationalmodules.
To meetall of thesegoalwill requiremanyyearsof programmer'stimeandlarger,faster
computers,in additionto thecontinuedevolutionof thetechnicaldesigncodes.A methodof
providing theinterfacebetweenuserandcomputerprogram,andbetweenmore thanone
computerprogram,canbedonethroughtheuseof anexpertsystemshell.TheRocketEn-
gineDesignExpertSystem(REDES)prototypeis aninitial attemptto achievethegoalsfor
betterrocketenginedesignthroughthesemeans.As well asencouragingandinsuringprop-
er utilizationof therocketengineevaluationcodes,thefirst stageof theexpertsystemde-
signwasto implementa methodof collectingdatafrom theuser,sharinginputsandoutputs
betweencomputercodes,anddisplayingtheresults.Theexpertsystembeingdesignedis
currentlyaddressingtheproblemsencounteredin trying to keeptrackof thelargeamountsof
datarequiredto run theavailableevaluationcodes.As thenumberof evaluationcodeswhich
arecontainedin theexpertsystemgrows,theamountof datarequiredto run themwill in-
crease,also.Thecapabilitiesof REDEScanbeexpandedto handlethesenewinputsin man-
ageablestepsandbeusedasadesigntool at eachstageof development.
EXPERT SYSTEM OVERVIEW
Theapproachtakenwasto constructtheexpertsystemwithin a shellin amodularfash-
ion.Eachpartof theexpertsystemwasinitially designedwith a low levelof complexitywith
the intentionof increasingthecomplexity(andfunctionality) in anincrementalfashion.For
example,theinitial analysisof thethrustchamberenergyreleaseis onedimensionalequilib-
rium (ODE)chemistry,but theultimatedevelopmentof thisportionof theprogramwould in-
cludestate-of-the-artfinite-ratereactionmodels.This moduleacceptedinputsfrom thesim-
plified injectordesignmoduleandprovidesoutputsthatarerequiredasinput to thethree-di-
mensionalregenerativecoolingchannelmodule.Oncethecomputationalloopis functioning
properlyfor thesesimplified modules,thecomplexityof anyoneof themodulescanbe in-
creasedwithout greatlyinterruptingtheflow of informationbetweenthem.Upgradingthe
f_
one-dimensional chemical reactions to include kinetic effects can be done later and only
slightly affects the data flow between modules. Similarly, the complexity of existing modules
can be increased or new modules can be added without having to make major changes to the
fundamental structure of the expert system. Proceeding from the simple scheme to the more
complicated can be done after other capabilities become operational and the interaction be-
tween parts is successfully demonstrated. In a similar manner, the functional complexity of
the injector element design, injector face design, regenerative cooling channel evaluation, and
the analysis of other parts of the engine, can all be initially integrated and tested at a low lev-
el of complexity.
The first "closed-loop" computation consisted of the injector design-energy release-re-
gen cooling computational iteration. Specific computer programs which were included in the
expert system include the Three Dimensional Kinetics (TDK) code 1, and the Rocket Thermal
Evaluation (RTE) code 2. Initial implementation of the TDK code only uses the ODE chemical
reaction program contained in TDK as a subprogram.
The expert system can accept input data describing a rocket engine, determine the re-
quired number and size of coaxial elements for that engine, determine the energy release of
the combusting propellants based on one dimensional equilibrium calculations, and calculate
the heat transfer to a coolant flowing through the regenerative cooling channels of the en-
gineusing a three-dimensional conduction code.
The structure of the data which define thrust chambers, nozzles, and other components of
a rocket engine was determined by input data requirements of the included computer codes.
The information which is collected for each component is only that which is required for the
computer programs which are currently available. Therefore, if pertinent data for component
#1 is A, B, C, and D, but code XYZ only needs A, B, and C, the input of D may not be provid-
ed for at the current time. In most cases, if the data is not needed, it is not asked for. Possi-
ble values for injector and propellant types were limited to facilitate the expert system devel-
opment. Injector design is limited to coaxial tube injectors and the only allowable types of fu-
el and oxidizer are diatomic hydrogen and diatomic oxygen.
The form of the user interface was created to provide ease of use and understanding
(figure 2). It consists of an "Engine Control" table (or panel) of parameters, divided into
groups (called subpanels, such as thrust chamber, injector, for example), and each group has
entries (also called slots) in the table where values can be input as well as table entries
where calculated output values are displayed. If a value is entered into an input slot which is
connected to a computation, the output value affected by the input slot will be updated imme-
diately. Some calculated values can be computed in many different ways. For example, the
thrust chamber diameter can be computed from either the injector face area or the contraction
ratio and the area of the throat. Only one method of calculation has been currently implement-
ed in the expert system. The flexibility of being able to calculate all possible quantities given
a limited set of inputs is currently being sacrificed to avoid infinite self-referencing calcula-
tions.
An input-output relationship exists between the user, the subpanels, and the computer
codes as shown in figure 3. The user provides inputs to the subpanels which compute output
values in other subpanels or provide input values for the computer codes. Interaction with the
program is through a menu selection or, in the case of numerical values, by directly typing
them in from the keyboard. The method of accessing data, creating datasets, executing a FOR-
TRAN program, and extracting the desired output data was implemented in REDES. When a
3
FORTRAN computer program is to be run, computer programs (called methods), v,Titten in
the computer language LISP, perform the functions of collecting the necessary data, create a
well-formed input dataset for the FORTRAN code, run the code, and finally, look through the
output for necessary information and then store that information in the expert system format.
This process is initiated by pushing a button (called a method actuator).
Computer codes ultimately provide the necessary data for conveying the computed infor-
mation back to the user (figure 3) in the form of graphs. Display of the program results in
graphical form was developed and extended to provide plots with multiple dependent vari-
ables. Figure 4 displays the panel on which the graphs are constructed. Method actuators
are used to plot nozzle contours or X-Y plots. Nozzle contours are specified in the nozzle
subpanel located on the Engine Control panel. Abscissa parameters and up to five ordinate
parameters are specified on the graphic panel via pop-up menu selection. The plotted param-
eters can be from the same computer run or from two (or more) different runs. This allows
comparative plots showing similar parameters from different engines or engine conditions.
SUMMARY
Designing a rocket engine requires the expertise of many disciplines which exist in both
the analytical and empirical domains. Computer codes presently attack a portion of the entire
problem at best, usually at the expense of user-friendliness or generality.
A prototype of a rocket engine design expert system has been designed using an expert
system shell to encourage and insure the proper use of a wide range of codes. The develop-
ment of the expert system is incremental starting with simple physical models and systems
and progressing toward completely detailed systems using transient, three-dimensional
state-of-the-art models. Energy release in the combustion chamber is presently modeled us-
ing ODE and the temperature distribution of the chamber wall is calculated with RTE. Current-
ly, the injector design is limited to coaxial tube injectors using diatomic oxygen as the oxidiz-
er and diatomic hydrogen as the fuel.
The expert system uses panels and subpanels to control input and output. Values which
describe the different pieces of the entire rocket engine are directly input from the keyboard.
Calculating or nondimensionalizing nozzle coordinates and imposing a minimum coaxial gap
size to determine the number of injector elements can be performed at the touch of a button.
Output values are displayed graphically in user-specified plots with up to five dependent
variables per plot.
The ultimate goal of being able to optimize an engine design considering all of the contrib-
uting, complex factors using a closed loop process is becoming a reality with the develop-
ment of the Rocket Engine Design Expert System.
REFERENCES
1. Nickerson, G.R., Coats, D.E., and Dang, L.D., "Two Dimensional Kinetic Reference
Computer Program, _K)" NASA CR-178628, 1985.
2. Naraghi, M.H.N., "RTE-A Computer Code For Three Dimensional Rocket Thermal
Evaluation" NAG3-759, 1988.
IP
4
INJECTOR
DESIGN*
REGENERATIVE COOLING
HEAT TRANSFER*
J
J
J
PERFORMANCE
ANALYSIS
INJECTION &
ATOMIZATION
COMBUSTION
INSTABILITY
NOZZLE
DESIGN
Figure 1. There are numerous technical disciplines which are involved in the design of a liquid rocket en-
gine. Those which appear with an asterisk (*) are currently included in the Rocket Engine De-
sign Expert System Prototype.
Figure2. On-screendisplayof the Engine Control Panel. Visible portion of metapanel is dark gray column
on the left and individual subpanels are to the right of each subpanel identifier display.
ORIGINAL PAGE
BLACK AND WI-CTE pFt_RAPH
SUBPANELS
COMPUTER
CODES
THRUST
THROAT
CHAMBER
ENGINE
INJECTOR
FUEL
OXIDIZER
REGEN
NOZZLE
ODE
RTE
USER
Figure 3. An input-output relationship exists between the user, the subpanels, and the computer
codes. Subpanels and computer codes both provide output back to the user.
Open Logo
Open ECP
Open GCP
Run ODE
Run RTE
Iil,tll!li,!lqllli lllllllt lm  li  l  ll,ltll,,l  11
Figure 4. On-screen display of the entire Graphic Control Panel showing nozzle contour of Current Engine
(ENGINE.ASE) and a specific impulse plot from an ODE evaluation of the Advanced Space En-
gine.
National Aeronauticsand Report Documentation Page
Space Administration
1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.
NASA TM-102373
5. Report Date4. Title and Subtitle
A Rocket Engine Design Expert System
7. Author(s)
Kenneth J. Davidian
9. Performing Organization Name and Address
National Aeronautics and Space Administration
Lewis Research Center
Cleveland, Ohio 44135-3191
12. Sponsoring Agency Name and Address
National Aeronautics and Space Administration
Washington, D.C. 20546-0001
6. Performing Organization Code
8. Performing Organization Report No.
E-5107
10. Work Unit No.
506-42-11
11. Contract or Grant No.
13. Type of Report and Period Covered
Technical Memorandum
14. Sponsoring Agency Coda
15. Supplementary Notes
Prepared for the 26th JANNAF Combustion Meeting, Pasadena, California, October 23-27, 1989.
16. Abstract
The overall structure and capabilities of an expert system designed to evaluate rocket engine performance are
described. The expert system incorporates a JANNAF standard reference computer code to determine rocket
engine performance and a state-of-the-art finite elements computer code to calculate the interactions between
propellant injection, energy release in the combustion chamber, and regenerative cooling heat transfer. Rule-of-
thumb heuristics have been incorporated for the hydrogen-oxygen coaxial injector design, including a minimum
gap size constraint on the total number of injector elements. One-dimensional equilibrium chemistry was
employed in the energy release analysis of the combustion chamber and three dimensional finite-difference
analysis of the regenerative cooling channels is used to calculate the pressure drop along the channels and the
coolant temperature as it exits the coolant circuit. Inputting values to describe the geometry and state properties
of the entire system is done directly from the computer keyboard. Graphical display of all output results from the
computer code analyses is facilitated by menu selection of up to five dependent variables per plot.
17. Key Words (Suggested by Author(s))
Rocket engine
Liquid propellant
Liquid rocket engine
Expert system
18. Distribution Statement
Unclassified - Unlimited
Subject Category 20
19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No of pages
Unclassified Unclassified 10
NASAFORM16=SOCTee *For sale by the National Technical Information Service, Springfield, Virginia 22161
22. Price*
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