A Transient Model of the RL 10A-3-3A Rocket Engine

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lfj i I -_.J NASA Contractor Report 195478 AIAA-95-2968

A Transient Model of the RL 10A-3-3A Rocket Engine

Michael E Binder NYMA, Inc. Engineering Services Division Brook Park, Ohio

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l_i National Aeronautics and Space Administration

A Transient Model of the RL10A-3-3A Rocket Engine

_tchael Bind_ NYMA, Inc. 2001 Aerospace Parkway Brook Park, Ohio 44129

Abstract

RL10A-3-3A rocket engines have served as the main tm3Pulsion system for Centaur upper stage vehicles since the early 1980's. This hydrogen/oxygen expander cycle engine continues to play a major role in the American launch industry. The Space Propulsion Technology Division at the NASA Lewis Research Center has created a computer model of the RL10 engine, based on detailed component analyses and available test data. This RL10 engine model can imxlict the perfcmnanoe of the engine over a wide range of operating conditions. The model may also be used to predict the effects of any proposed design changes and anticipated failure scenarios. In this paper, the results of the component analyses are discussed. Simulation results from the new system model are compared with engine test and flight data, including the start and shut-down transient clmtaoeristics.

1.0 state performance of the RL10A-3-3A, but the predicted lime required for the engine to reach a specified thrust The RL10A rocket engine is an important component during engine start (time-to-accelerate) showed of the United States space infraslntctme. Two RL10 significant differences with test data 2. It is believed engines form the main propulsion system for the that these discrepancies were due to errors in Centaur upper stage vehicle, which boosts commercial, exuapolating the available component perfmmance data scientific, and military payloads from a high altitude to cover engine-start conditions, as well as errors in the into Earth orbit and beyond (planetary missions). The physical models used for heat umsfer and two-phase Centaur upper stage is used on both Atlas and Titan flow. Analysis of each RL10 engine component was launch vehicles. The initial RL10A-1 was developed in undertaken in order to verify the origin of the dam the 1960's by Pratt & Whitney (P&W), under contract lXovided by P&W, and to improve the accuracy of the to NASA. The RL10A-3-3A, RL10A-4, and RL10A- models at far off-design conditions. These analyses 4-1 engines used today incorporate component were also used to benchmark our ability to acowalely improvements but have the same basic configuration as model new rocket engine designs forwhich test data are that of the original RL10A-I engine. RL10's have not yet available; the RLIO engine system provided test been highly reliable servants of America's space data to validate the available component and system program for over 30 years. The RLIOA-3-3A engine is modeling tools. represented schematically in Figure 1. In this paper, the RLIOA-3-3A rocket engine and its The Space Propulsion Technology Division (SPTD) at various components me described briefly. The analysis the NASA Lewis Research Center began developing a n:mlts for each component are then discussed, including computer model of the RL10 in 1991. This model was comparisons with existing component test data. The intended for government use in engine system research, new engine system model, which includes the results of _-analysis and flight failure investigations. The selected component analyses, is described and first version of the model was created using data pn_dictionsof me model_ mmparedto gn_-test and provided by Pratt & Whitney, and the ROCket Engine flight data. For a more detailed discussion of the Transient Simulator (ROCETS) I system analysis modeling work summariz_ here, the reader is referredm program. This model could accurately p_llct the steady- theRL10A-3-3ARocketEngineModelingProject conditions exlgdenced by the pom_ during engine tort F'mMReport3. and shutdown. P&W had also wovided the SPTD with test data maps of turbine efficiency and flow zesistance As the simulation results will show, the new RL10 as fanctlom of overall pressure ratio sad velocity ratio model ccuectly predicts variation in engine transient (u/Co). These maps do cover a range of conditions behavior due to inlet conditions, initial thermal suitable for engine start and shutdown sinudatious. conditioning, and ignition delay. 3.2 Detailed Pumn Analyses 2.0 RL10A-3-3A Engine Description Two different analysis codes, PUMPA 4 and LSISO 5, were used to model the RL10A-3-3A fuel and LOX The RL10 engine design (all models) is based oll a full pumps. Tan pamp head coefficients _ by each expander cycle, as shown in Figure I. Hydrogm fuel is code agree with test data to within five percent (5%) used to cool the thrust chamber and nozzle, and the over the engine's normal steady-state operating range. thermal eaezgy Wansfetmi to the coolant is used to drive The PUIvIPA and LSISO efficiency predictions, the aubopumps. Warm hydrogen gas is injected with however, differed from test data by as much as fifteen cryogenic liquid oxygen into the comlmslkm chamber percent (15%), and could therefme not be used in the and burned to provide thrust. During enghg sum, fuel engine system model. PUMPA was also used to tank iaessme and the initial ambient heat in the cooling predict the performance of the RL10A-3-3A pumps at jacket metal are used to start rotation on the mrbiue. the engine start conditions. The results of these After ignition, the heat of combustion is used to analyses were used qualitatively to help extrapolate the accelerate the tmbopumps to full power. Because the head maps beyond the available test data provided, as Centaur upper stage vehicle uses two RL10 engines, it discussed later in this section. It sJz3uldbe noted that a is important that the engines start simultaneously (to subsequent version of the PUMPA code was recently minimize thrust imbalances). For the purpose of developed which better wedicts the RL10A-3-3A pump providing a quantitative measure of the engine start design point efficiency, without affecting the head times, we shall refer to the time between the start predictions. The new version of PUMPA was signal and the chamber pressure reaching 200 psla as completed too late to allow a comprehensive a_tlysls of the t/me_/erate. start coalitions to be perforated again for this project.

During engine shutdown, the fuel inlet, fuel shut-off, In addition to the PUMPA and LSISO analyses and oxidizer inlet valves are clmed. The oumbustion above, a third analysis was peffcmned which process stops mul the fuel and oxidizer drain flora the was spec_xcally designed to estimate the lOW speed engine system; LOX drains out through the thrust pump head (as experienced during start). This method chamber, and the fuel drains out through the pump was suggested by Rostafmski 6 and requires that the cool-down valves. design point perfommuce of the pump be known. This when cambiued with a separate model of the 3.0 Turbomachinerv Analysis pump exit diffuser, appears to match well with the limited test data available at engine start conditions. 3.1 Turbonumu Background Information Although Im3mislng, this modeling technique proved The RL10A-3-3A Uubopump includes a two-stage fuel impractical for transient system simulation (slow turbine which drives a two-stage fuel pump ou a execution, numerical instabilities, etc.) and was shaft, rex! a single-stage LOX pump tla-ough a therefcce not included in the new RLI0 system model gear box. At the engine's normal operating point, a fuel flow of 6 lb/sec is pumped to a wessure of 1100 Using available engine test data and informatim gained psla, and 30 Ib/sec of LOX is ptmapedto 600 psia. The from the analyses discussed above, the pump normal operating speed of the fuel pump is 32000 rpm, performance maps provided by Pratt & Whitney were and the LOX pump speed is 12800 rpm. exlrapolaled to include _nditions at engine start and shutdown (zero speed, zeverse flow, ca_ etc.). In Pratt & Whimey provided the NASA SPTD with test order to represeut such a wide range of operating data maps of head coefficient and efficiency for each conditions, a map format suggested by Chaudl_y 7 is pump stage as functions of flow coefficient, and used. The new map format defines normalized head included a speed correction factor f_" efficiency. These parameter(h) and torque parameter (13)as functions of a maps do not cover the entire range of operating third parameter, 6 (them) as described below. The new pumpperfmmancemapsfortheengine system model output and test data have not been explored; the me shown in Figures 2 and 3. TURBA code is still considered to be in the development phase. The performance maps provided by P&W have flmrefore been retained in the new system model. h= I_= N Q + N The turbine analysis performed in this study indicates that it is possible to estimate the design point performance of a new turbine to within a few percenL It is also possible to predict the overall trends in perfmmance at off-design conditions. As with the pump analyses discussed above, however, the accmacy of the turbine lm_iictious may not be sufficient for use in transient or deep-throuling simulations of a new engine. When component test data is available for a whe_ ffihead (in feet) new turbine design, it might be possible to adequately N = shaftspeed (inrpm) adjust the model based on only a few test data points. Q = volmnetricflow(ingpm) andthesubscriptd denotesthedesign 4.0 Thrust Chamber and Cooline Jacket condition. 4.1 Thrust Chamber Backtwound Information walls of the RL10A-3-3A thrust chamber are The results of the pump analyses descn'bed above indicate that it shouki be possible to predia the general consmgted of stainless steel tubing. Hydrogen fuel is walls performance characteristics of new pump designs. pumped throegh these robes in order to cool the Results from this type of analysis are valuable for of the thrust dmmber and provide thermal en_gy to the conceptual engine design and simulation activities. turbine. The robes are brazed together and reinforced Such component Ixedictions may not be sufficiently with bands on the outside, as well as a metal girdle accurate for use in engine start-transient simulations, around the throat sectim. A silver throat insert is cast especially if no test data is available with which to in place to increase the nozzle area ratio and specific anchor the new pump models. impulse. The thrust chamber normally operates at a pressure of 475 psi& a mixture ratio (O/F) around 5.0, 3.3 Detailed Analysis of Fuel Turbin_ a thrust of 16500 lbf, and a specific impulse of 445 The RLIOA-3-3A turbinewas alsoanalyzedruingthe seconds. TURBA code s, which is cmrently being developed at the NASA Lewis Research Center. TURBA is a one- The analysis of the RLIOA-3-3A thrust chamb_ was dimensional mean-line code which combines basic divided into three basic areas: 1) cooling jacket heat physics (vdocity trianglesand isentropicrelaX'ms) transfer, 2) combustion chamber performance, and 3) with empirical cot_lations derived from existing nozzle performance. Each analysis is described below. uubine designs. The turbine perfmamnee predictious could notbe directly compared with the maps provided 4.2l_etailfd Analysis of Coolint Jacket Heat by P&W. Instead, both sets of maps were used as T_mrer inputs to a simple turbine simulation, and the resulting original model of the RL10 cooling jacket had only overall efficiencies and flow rates were compared. five heat transfer cakulatm nodes distributed along the Although the overall performance trends i_edicted by cooling circuit This model was considered to be too 1XJRBA are similar to those indicated by the P&W coarse and amore detailed model was c_mted for this data, a more quantitative comparison shows that project. significant differences exist. The predicted overall turbine effgietgy, for ex_m_ diffe¢_by more than5% CET93, a _ equih'laium program 9, was ffmn the P&W data, especially at low speeds. It has used to refine the table of hot-gas properties. The been fmlh_ noted that relatively small variations in the Rocket Thmnal Evaluator (RTE) code lo was used to turbineperfotmaige at low speeds can profoundly affect predict the flow resistance of the cooling jacket and the the RL10 engine time-to-accelerate. Possible effects of tube curvatt_e ou heat transfer rate. Heat explanations for the poor match between TURBA transferbetween the combustion gas and chamber walls waspnxk_ ustagan euthalpy-driven Banz correlatim 4.3Detailed Analysis of Combustion I t. The euthalpy gradient was used instead of the Chamber Performance temperature gradient because this more accurately In addition to revising the ¢ombmtion gas property predictsvariationinheatIransferatdifferentmixture tables f(g the new model, several other imwovements ratios. A Colbum correlatioa 12was used to detetlBi_ were made. In the original RL10 engine model, the the heat transfer from the chamber wall to the coolant thrust chamber was treated as a _ingle point, without flow. It was clisc_vered that cembining twenty hot-gas considering axial variation. In reality, there me and metal property nodes with five (rather than 20) m,wnmmwn losses and eJ_ager in static pressure along coolant nodes couM significantly increase the the hot-gas flow path which will affect performance. computational efficiency of the transient system model These effects were relatively simple to wedict and were without loss of accuracy overall. This the added to the model. Au analysis of the RL10A-3-3A configurmkm was used in the new RL10A-3-3A system thrust chamb_ _bly was also perfonned usingthe model. ROCCID code 13, which provides a c_pability of modeling the propellant injectors, atomization and Figure 4 shows the predicted heat flux, wall combustion processes. The objectives of this analysis temlmauue, coolant _ and wesm_ along the were to validate onr _ty to wedict ¢*_ axial length of the thrust chamber cooling jacket. Test usingRL10 data from P&W, and to extend the range of data show_ ux_ _ ta tempenm_ and _ mixture ratios represented in that data set. The RL10 are not available for comparismx. The accma_ of the injextef proved difficult to model using ROCCID; new heat Wausfer model can only be judged by the several aspects of its design are not found in the more overall _ rise and wesuue drop across the contemporary designs which ROCCID was intended to cooling jacket. Based on these parameters, an model. As a result, the results of the ROCCID empirical _ of 1.08 was added for the hot-gas analyses did not show a good match with the P&W heat transfer coefficient and a fact_ of 0.94 applied to data. 2_e R_ model also _ numea'ieal the predicted jacket flow resistanoe. These empirical convergence woblems at low pressures (below 160 correction factors represent average values, since the psla), where the c*-eff'tciency changes significantly. actual heat _ metlicient _ to vary somewhat Tbe data maps wovided by P&W have beea retained in fiem one RLI0 engine to another. These variations the new system model may be due to small mmmfactming diHe_aces; they are not c_nsidered critical as loag as the engine has During the engine start sequmce, heat transfeg in the sufficient starting pow_. injector can play a discernal_ role in the system's dynamic behavior, prima_y by clumgingthe densityof A simple cme-dimeasiem_ film boiling model was also the injected LOX. Simple models of heat ttm_er in added to the oooling jacket heat trm_fer model. Fdm- theinjectorekmems endinter-_t bulkheadwere boiling effects have been suggested as the cause of the added to the new engine model. Although the_ is four to eight Hertz pressure oscillations often insufficienttestdatato validate tbe modeis, the results ex_ dsmg the RLI0 engine startseqne_e. Tee al_earmasmable.The additionofthesemodeledeffects new model still does not show these pressure delaysthetime-to-accelerateby apwoximately I00 oscUlations; they may be due to two-dimensional milliseconds,Considered over all engine start effects not modeled here or to local choking within the transients simulated, this delay results in a more two-phase fluid. accurate predktion of time-to-accelerate. Figure 5 shows the _ heat tnmsfef rate in the injoc/_ as a The analyses wes_ted hem demmsuale the capab_ty function of time during a typical engine start. of one-dimensional models to pmdi_ the effects of various oondifims on heat transfer. Depending on the 4.4Detailed Analvsta of No_,.zle Performance accuracy required for system simulations, some The RLIOA-3-3A nozzle perf_ affects the adjustment to the heat transfer _ts using test mmbustion chamber Wesmre and flowrate, as well as data my be required. Test results ate also useful in the specific impulse and _-ust of the engine. Prau& defining the variability in heat trm_e_ characteristics Whitney had originally ixovided nozzle perfm_ance due to manufacturing tolerances and other factors. data in the form of specific impulse (Isp) tables with additional corre_ons for various kinetic losses. Analyses were performed at Lewis using a Two D:on_l$ion_l _es (TDK) pl'ogram 14in O_ to better understand the P&W data. Figure 6 shows the upstream of the orifice or valve. This modeling output of the TDK analysis compared with the P&W approach was used fog the LOX injector elements and data. The results match well at the engine's normal fuel cool-down valves. Two-phase flow in the oxidizer operating point of 475 psia and O/F = 5.0. The control valve is modeled as incemweuibie, limited by wedicted and P&W values differ mote significantly at the saturation pressme of the fluid until the flow low pressures and mixture-ratios, however. The becomes entirely gaseous, after which it is Ireated as predicted Isp maps have been included in the new isentropie flow of an ideal gas. These models agree RL10A-3-3A system model. well with avaflabie test data.

Several different _ were takm to detmaine the The fluid resistances of ducts and tubes are typically nozzie discharge o3efficknt (CA). P&W had specified a determined by flow testing those components. For CA of approximately 0.98. A Navief-Stokes analysis 15 new rocket engine systems, empirical data of was performed at Lewis which wedicted the discharge this kind may not be available during the analysis coefficient to be 0.979, a remarkable agreement. phase. A simple ol_ff-dil3_e_onal analysis19 of flow in Trimming the CA value used in the engine simulation an RLI0 duct (from the turbine discharge to the main to match lxedicted chamber laesmre with test data gave fuel shut-off valve) was peffmmed and the results wea'e a value of 0.975. The TDK analysis descn3_ above compared with the resistmwe _ by P&W. Tne had further indicated that the CA may ch_mge somewhat inflate roughness on the interior of the duct was not with chamber pressure and mixture ratio. After known, so we considered a range of options from considering these various results, a constant CA of smooth commercial steel pipe to drawn tubing. The 0.975 was selected fog use in the new system model. analyses indicated a range of possible K valnes19 from 0.928 to 0.487; the value of g given by P&W was $.0 Miscellaneous Comnonents 0.648. Our estimates the_fore define a range of possible values which bracket the suggested value with In general, the ducts, valves, and manifolds in the RL10 m each" of 25 to 43 %. The _ce provided by engine were not analyzed in detail Many of these P&W has been retained for the new RL10 engine components have complex geometries that would model, but this analysis suggests that we c_mwobably require f'mite-element methods to model im31gdy. In estimate the resistance for a new (untested) duct to the case of the fuel pump cool-down valves, oxidizer within 4./- 30%. Better estimates might be possible if control valve, madLOX injectar elements, however, the the surface roughness of the intended duct is well models for two-phase flow contained in the original defined. model required improvement. We also attempted to verify the resistance of a single duct as specified by It is evident from the discussion above that accurate Pratt & Whitney using generic one-dimensional one-dimmsioml models of ducts and valves in a new metho&. engine design will require at least some flow testing. Befme such data is available, it would be prudent to During the engine start, several components experience _tsider the effects of uncertainty in engine system two-phase critical and mrJmked flow conditions. The simS. In the case of valves and ducts where two- fnel-pump cool-down valves, which vent liquid phase flow might exist, it is advisable to test the hydrogen overbom_ ale always clinked and thehydrogen components over d_Jr entire operating range, since flashes to vapor as it is vented. The oxidizer control two-phase effects caa often lead to unexpec_ behavi_. valve and LOX injector elemmts experience two-phase Flow models which include two and three dimensional flow for only a small period of time during start, effects may also woduce more accurate resistance transitionin8 at some point between choked and predictions. unchoked conditions. The challenge was to devise models which allow a relatively continuous transition 6.0 _lew RL10A-3-3A En2ine System Model between the various flow conditions during start. Tne new RLIOA-3-3A engine system model includes A number of different _ were cousidered 26 J7 the results of several of the detailed component 18. Ultimately, a model was derived which treats the analyses, as described above. In addition to these flow as incompressible, but limits the assumed comixagnt model changes, several improvements wae dowusueam wessure to either saturation og isenlropic made in the structnre of the system model itself. critical pw._sme, depending on the value of the pressure Tracking of the total-to-static conversions for pressure and mthalpy was iml_ved in the new system model, is unl/kely that the values were all Wecisely 20 Ibf- for example. in.

It became necessary to create two difterent models of 6.1.2 Uncertainties in Valve movement the RL10 engine: one for ¢imzdatlng Start transient "Fne transient behavior of the engine in both start behavior and steady-state peff_ and the other for and shutdown is largely determined by the opening simulating shut-down transient behavior. During shut- and closing of valves. Variations in valve and down, the ducts and manifolds in tbe engine me emptied actuator behavior actually do occm for a variety of into space, and dynamk: volumes had to be added to the reasons. In some cases, the opening and closing model to allow the simulation of these effects. times of valves can be infen'ed f:n:an test data. In Including these dynamic volumes in the start transient most cases,however, this is not posm'ble because and steady-state model changed the predicted start of the limited number of engine sensors and their transient behavior significantly, in disagreement with dynamic response rates. Valve data provided by test dam. These differences could not be resolved, and P&W has been cmnbined with information infm'ed so two separ_ modeLs w_e developed. from avaUable test data to define 'typical" valve movement schedules for the new system model. 6.1Effects of Modelin_ Uncertainty This single set of typical scigdules was used for all Before discussing the output of the new system model, simulations performed in this study. it is important to note several unresolved sources of uncertainty in the model which will affect our ability to 6.1.3 Uncertain Initial conditions accurately simulate a given RLI0 engine firing. Tlw.se The temlga'atme of the combustion chamber, uncertainties can be divided into four categories: 1) nozzle and cooling jacket at the beginning of the enceminty in hardware _ 2) _ties engine start sequence is an importmt factor in the in valve dynamic behavior, 3) _ties in engine engine timc-to-acccler_. Unlike the engine inlet initial conditions, and 4) uncertainty in the main and tempetatm_, there is no foible chamber ignition delay. _ is also a great deal of meamtemmt of initial jacket _ for any non-linear interaction between RL10 engine given test or flight. Temperatures that are components 2o. Charactaizing the _,nction between measured on the engine genea_ily show false the various operating lXtmneters with uncertainty was readings before start due to ambient conditiot_, beyond the u:ope of this study. metal _ with other c_a_ and the lack of pmpeUant flow at that time. Tha initial 6.1.1 Uncertainti_es in Hardware temperature of the coolingjacket, ducts, manifokis, eh_wtt,ri_tiot and ether components must be estimated, often There LSsome mceminty in the actml value of the based on limited information from past testing. discharge coefficient for the fuel-immp cool-down valves. In the RL10 model, the discharge In the RLI0 model diacussed here, tbe_ coefficimt is set at 0.6 for grom_test and 0.8 for of the cooling jacket is asmanecl to be a uniform flight. These values were chosen based on 540 R for furst butas and 350 R for second bums. discussims with enginec_ at Pratt & Whitney but Tha _oling jacket ialet manifoM is asmmod to be no real caWntion data is available to verify these at 200 R became the inlet manifokl is extmm_ to values, l_e res/stmce of the cool-down valves is some of the fnel pump toM-down flow before start. .n important factor in the engine time-go-accelwate. All other _ in the systan m'e assumed to be in thermalequflflx-imnwith thewopellantflows '1"ne drag tmque (due to bearings, seats, gears, etc.) at start. Because these assumptions are somewhat of the RLIO Imbopump (fuel and LOX combined) arbitrary, fl_ey me h'kely to be in avor to some is a known f_wce of mgine-to-engine variation. degree for my given firing. Tbe valne LSnot genially mea._d for ea,'h engine but past studies have shown that the torques v&y 6.1.4 Uncertainty in hmition delay from 8 to 36 Ibf-in (with resla_::t to tbe fnel pump For the simulations considered here, the ignition shaft)2o.A constantnominal value of 20 lbf-in times were set mammlly to agxee with the meamnxl has been used for all simulaficms nm for this study. data. In order to simulate m engine start for which It is uncertain what the actual values of running data is not yet available, a model of the ignition torque were for the lest and flights considered but it Igocess would be required. 131is model could be based on theorelkal analysis, or might be derived conditions. The new system model's steady-state from test dau_ NASA does not currently have an are therefore slightly less acowale than those ignition model for the RL10. of the original system models. It was decided, however, that the turbine maps suggested for start transient 6.2RL10 Steady-state Engine Performance modeling would be used throughout, and the associated steady-state mor accepted. Ten test cases are considered for the steady-_aate performance predictions. Five tests are based on 6.3RLI0 Start Transient Simulations diffe_ntquiescentoperatingpointsf_ a single ground- The results of RLI0 start transient simulations were testrunof a singleengine(EaglneP2087,Run 2.01, compared with both ground-test and flight data. Figure Ocloi_ 4,1991). The other five tests are based on the 8 (a - e) shows the predicted and measured start final states of five sturt-Umsient data sets (five different Wamients of a single ground-test first-burn. Figure 9 ground-testruns) of a singleengine(P2093). Flight shows chamber pressure and pump speed data for an data has not been included in this comparison because Atlas/Centaur flight (AC-72), while Figure 10 shows insufficient data exists to determine the mixture-ratio similar data for the second burn (restart) of a different and trim position of the oxidizer control valve (OCV) flight (AC-74). In each of these runs, the ignition time for those firings. For the ground-test runs considered, has been setin the model based on examination of the the OCV position has been trimmed in the simulation testor flightdam. 1_e differencebetweengronnd-test to achieve the steady-stale mixture ratio indicaledby the and flight engine simulations is the value chosen for test data. Since the OCV position is not a measured the fuel cool-down valves discharge coefficient (which parameter, the simulated trim position could not be reflects differences between the vehicle and test-stand verified directly with test data. A comprehensive ductwork). The difference between first and second bcm performanee wediction for a typical case b shown in simulations b the a_umcd initial temperatm_ of the Table 1. In general, only a few parameters are actually combustion chamber metal. These assunwM variations measmed on engine firings (14 parameters on grmmd- wornalso discnssed in section 6.1 above. tests, 8 in flight). Of the fotmeen pmmneters measured in ground-tests, five are used as inputs to the model The start model generally matches the measured tim¢- (inlet conditions and chamber lZessure), and so only to-accelmme of the engine to within approximately 230 nine _ _ are cmapm_ with test data for milliseconds, using only estimates for initial each case. _ bearing fxicti_ valve u:heduies and other factors which may vary from nm to run and from Figure 7 shows the distribution of error between the engine to engine. Table 2 gives the predicted vs. measured and predicted parameter values in the ten measured finw-to-a_x_ate for six ground-test and three ground-test cases. The model lX'edictions match the flight-engine firings. One of the flight simulations is meastm_ values to within 10% fcf all pmmnetefs on all off by 280 msec (rather than 230 reset), but this tests (a total of 90 values). Most wedictions are within appears to be an aberration relative to other flight- 3% of the test resuRs. The most significant errms are engine starts. Comparing the results of this start in the turbine inlet temperature and the pump discharge Uan._ent with those from other frights, it appears likely _esmres. The difference between the wedicted and that the conditions f_ this flight were differeat in ways measuredturbineinlettemperaturevariesfrem engine oth_ than their inlet conditkms alone. The model to engine,as discussedin section4.2of thispaper. correctly predicts start variations due to different engine The en_ in the pump discharge ptesmres appems to be inlet conditions, initial thermal conditions, and associated with mfl)cpump speeds that are consistently diffeamces betwem gnmnd and flight hardw_. lower than measm_. This _ in speed is mnst likely due to small errors in the turbine maps and The reader may note from Figures 8-10 that there are cooling jacket model; these errors cannot be easily some transient differences between the predicted and corwxted for without adversely affecting the wedicted measured chamber pressures which occur after the stm behavior. The turbine performance maps wovided engine bootstraps but before it reaches the quiescent by P&W for transient simulation are not the same as state. The small oscillations evident in the test data me those originally provided for use in the steady-state due to oscillations of the Thrust Control Valve CI'CV) model. The original maps do not work well in se_vo-mechanisnL The simulation does not include a simulating the start transient but the new maps do not model of the actuator dynamics, but the TCV is match as well at the engine's design operating assumed to open as a simple linear function of combustionchamberpressure. The simulation 7.0 Concludina Remarks thereforeoverMmotsthedesiredchamberpressureand The major goals set for this pmjeot were to create a doesnotoscillate,htseveralcases,tbesimuiatimdoes tramient model of the RLIOA-3-3A rcdwA eagine for showsceneunusualuausiemsbeforereacbtagsteady- government use, to betu_r understand the engine and its state;theseapsgartobedueto volume dynamics tn the mmlmnems, m_dto beaclanmk tha available cmnponem LOX pe_ iuiet ducc As the OCV suddmly opens and malysis tools using an existing mgine design. These the LOX system _ the simulatim pmlkts goershavebernacmmpUstet o,cmaems musodby nutd inerea, and phase changm. Thesemmsients,which Ke not evident Tha new RL10 start trmsieat model accmm_y Igedicts in the test data, may occtw in the simulations becanse tbeenginet e-t , me whm to ground- OCT serve dynamics m'e not included in the model. test and flight data. The model can granulate engine Thesetramientdifferencesbetweenpredictkmand test start transients over a wide range of inlet conditions, ate not considered significant; they would be minimized initial themml conditious, and ignition delays. This if models af tbe TCV and OCV actuatms are deveicj_ model also paedicts steady-state ped'ommance values in the futm_. which are within 10% of the meamred values in all cases, and within 3% for most _. The new To demmsuate one potmtial applkation of the system RLI0 shutdown model successfully reproduces the start model, Figure 11 shows the predicted metal eDgi_'s transieat behavior after main eogine cut-off. of the combustion chamber just upstream These new system models muld be used in the future to of the thront (its hottest point). This _ is not predict the effects of cJumges in the mgine design, and measm-ed, even in ground tests. The temperatme in this to simulate off-nmninal opmuing conditious. case peaks at amend 1875 R, which is a few hundw_ degrees below the melting point of the silver throat i,!performingthedmanedcomponentanalysesdescribed insert. Infoanation of this kind can be used to help in this patsy, a great deal has been leamod about the determine conqmtzmt wear and to _ the impaa of RLI0 e_gine. This activity has also provided am _ or hmdwa_ chaqes to tbe engine. oplmaunity to compare the output from available component modeling tools with test data from an 6.4RLIO Shutdown Transient Simulations existing engine design. Comparison of the Two firings have been used for comparison between mmlysisresalts with data provided by Pratt k Whitney model predictions and measured data. RLI0 eagine indicates that at least some empirical correction must be shutdowns do not appear to have any distinct feature made to the results of the component models. Such analogous to the time-to-accelerate for start transients. component models are nonetheless valuable in Although there are subtle variations in the rate of predicting the off-design _ of the engine deceleration, the nature of these differences is not as components, especially once _ cozre_ons have well understood as in the case of engine start. been included. Detailed three-dimensional computational-fluid-dynamic models may also be Figure 12 (a-d) shows the wedicted vs. measured considered in the future to improve the _tncy of shutdown for a gronnd-test engine. The RLIO mmponentperformancept fictious,thoughevenugh shutdown model has capturedmany interesting effects advancedtechniqueswillinvolvesome uncertainty, that occur during shutdown. In Fignre 12c, for especiallyfornew cx_aponentde&igus,The capability example, the simulated and measmed venturi pressmes may notyetexisttoweci_y predictthebehaviorof both show a characteristic dip, rise and then falioff in new componmts of engines for which no test data is the fuel venturi upstream pressure. This feature is available. In inch cases, the best that can be expec_d caused by the dynamic interaction of the fuel pump is to define a range of performance and transient cool-down valve opening gad main fuel shutoff valve behavior based on malysis. This type of infomation closing. Anotber _g charactedstic of Ibe RLI0 can be extremely valuable in the design and shutdown transient (as shown in Figure 12c) is the development of new ccmpommts or systems, especially jump in fuel pump inlet wesstwe due to reverse flow in combination with probablistic and uncertainty through the fuel pump. analysis techniques. Acknowledgements References

'Ibiswork was Ix_fonned undexconlractNAS3-27186. 1. Pratt & Whitney Government Engines, System Design Specification for the The anth0f would like to acknowledge the participation ROCETS System - Final Report, NASA of the following individuals, who performed the CR-184099, July 1990. component analyses which were incorporated into the new RL10A-3-3A system model. 2. Binder, M. - Sverdrup Technology Inc., Tom Tomsik (NASA LeRC) - Heat Transf_ An RLIOA.3.$A Rocket Engine Model Joseph Veres (NASA LeRC) - Turbine and Pumps Using the ROCETS Software, AIAA Ken Kacynski (NASA LeRC) - Combu.qion and Nozzle Paper 93-2357, June 1993. Doug Rapp (Sverdrup / NYMA) - Hot gas properties Dean Scheer (Sveldrup / NYMA) - Pumps 3. Binder, M., Tomslk,T, Veres,J, Albert Pavli (Sveadmp / NYMA) - Inject_ RLIOA-3-3A Modeling Project- Final William Tabata (NASA LeRC) - General RL10 Report, NASA TIM (number pending), information and expert_. 1995.

Pratt & Whitney (West Palm Beach, Florida) also 4. Veres, J, A Pump Meanline Analysis participated in this work, supplying us with design Code (PUMPA), NASA TM.106745, infoflnation and test data forvalidation. October 1994.

$. Gulbrandsen, N. Centrifugal Pump Loss lsolatiou Program (LSISO), COSMIC Program # MFS-13029, April 1967. *(note that the version of LSISO used in this study is not publicly available, this reference gives a general description of the program in its original, publicly disseminated form).

6. Rostafinsld, W, An Aulytical Method for Pr#dictiag the Performcmce of Centrifugal Pumps During Pressurized Startup, NASA TN D-4967, January 1969.

7. Chaudhry,H, Applied Hydraulic Transients - 2ml Edition, Van Nostrand Reinhold, New York 1987.

8. Veres, J., A Turbine Meanliae Analysis Code (TURBA), NASA TM (number pending), 1995.

9. McBride,B. and Gordon,S. CETg$ and CETPC: An Interim Updated Version of the NASA Lewis Computer Program for Calculating Complex Chemical Equilibria witk Applications , NASA TM-4557, March 1994. 10. Nnragbi, M.H.N, RTE - A Computer Code for Tkree-Dimensiolud Rocket 20.RLI0 Start Capability Working Group Thermal Evaluation, (NASA Grant NAG (government/industry cooperative), AC- 3.892), July 1991. 71 Failure lnvestigatiou - Final Report, (no contract or report number given), 11.Bnrt_ D.R, Smrvey of Relationships December 1993. between Tkeory aml Experiment for Convection Heat Transfer in Rocket Combustion Gases, Advances in Rocket Propulsion, AGARD, Technivision Services, Manchester, England, 1968.

12. Holman,J.P, Heat Transfer - Fourth Editiom McGraw Hill, 1976

13.Muss, J. and Nguyen,T_ User's Mauual for Rocket Combustor Interactive Design (ROCCID) arid Aualysis Computer Program, NASA CR-1087109, May 1991.

14. Nickerson,G, Two-Dimeasioual Kinetics (TDg) Nozzle Performauce Computer Program. Users Manual, (NASA Contract NAS8-39048), March 1993.

15. Kacynski, K., Calculut/ou of Propulsive Nozzle Flowfieid in Multidiffusing, Ckemicallyreacting Environments, NASA TM 106532, 1994.

16. Henry, R. and Fauske, H, Tkc Two- phase Critical Flow of Oue-Compouent Mixtures in NooJes, Orifices, amd Short Tubes, ASME Journal of Heat Transfer, May 1971.

17. Applied Physics Laboratory, JANNAF Rocket Engine Performamce Test Data Acquisitiou aml laterpretattou Manual, CPIA Publication 245, April 1975.

18.Shnonean, R. and Hendricks, R., Generalized Clwzrts for the Computation of Two-phase Choked Flow of Simple Cryogen Fluids, Cryogenics, February 1977.

19. Crane Co. and ABZ Inc., Crane Compauiou to Flow of Fluids Through Valves, Fittings, and Pipe, (software), 1992.

10 Oxkber Coatm/ Vdve (OO0

MBind_ / M.Milis

Figure 1 RLIOA-3-3A Engine System Schematic

1! 1.50

1.00

"'m

m l, _ L 0.50

-'I

I. / 0.00 0 E m L, 'P -0.50 ii FP Ist Stage o :]:: FP 2rid Stage -1.00 ...... Ox P.mp

-1.50

-2.00

0.00 0.50 1.00 1.50 2.00 2.50 3.00 3.50 Speed/FlowParameter (the,-)

Figure 2 Extrapolated Pump Head Maps for RL10 Model

0.70

0.60 0.50 / ! °..°,,,. .--- I °''° II/ --...... I 0.20

o.lo

o.oo l:P 2rid Stake

.o.lo ...... Ox P.mp _

-0.20

0.00 0.50 1.00 1.50 2.00 2.50 3.00 3.50

Speed/Flow Parameter (theta)

Figure 3 Extrapolated Pump Torque Maps for RLI0 Model

12 ]h,_dl,elod lleel Plu deq Jmelml ]P_41mmlIiml Taqmnm_ mlmqau_m, Wdl

2O J 2000 ,_ ! 18 _844 It._ 14 .1440 r e,I V i, f200 -\ J il 1444 !%+. lO 1000 I - !' •-J IX 404 | I\ 4 i % _ 404 2 2OO I m 4 4

-24 -14 0 14 24 60 4O SO -20 -14 4 14 24 64 44 SO

a.Vd I_dJm - Hed. mldpdmm (im+b_ _ 11n_

FT_lhmd CLdml T_NSMIT _lll J_

S44

l 1464 4S0

K. 0 I08O 4OO /

644 _ 464

| _54 J,, "% tl 200 / 1S0

140 / SO i-TJ ° 4 764 -24 -1O 4 14 20 50 44 SO *s+O -14 4 '14 14 54 44 60

J_d r'd'_" • Ned. ladpdmb (kud._ m_ _) &dd e'ud_. Ne.dt.]r,Jdtm_ C.udn. -'n 'n,n._

Figure 4 Predicted Axial Variation in Heat Transfer with New Model

]3 8O .07O // I /

a_ 4o

J lOor

o.oo 0.SO 1.oo I .SO 2.o0 2.So s.oo s.S0 4.00 Time from MES (mc)

Figure $ Predicted Heat Transfer in Injector Piena during Engine Start

45O

4OO

350 L J x/l.:

300 II_ib ° %

"1% 25O ---- - NtWIqlmWdi_-lS m I! n ---4--- PaW mSllmml _ 2OO

0 2 4 8 8 10 ]m_ st_, (o_

Figure 6 TDK/ODE Predictions of Isp vs. P&W Suggested Values

]4 0

Is 100%

90% 0 I-, 80% "_ 70%

--a 20% :l E= 10%

Percent Error relative to Measured Values

Figure 7 Steady-state Predictions vs. Measurements from Ground-Test Engine Runs (all runs, all parameters)

]6 8OO 14400

14000 SO0

20OO 400 J J tO000 // SO0 ? 10@0 J lO0 i 8000 i I r_ I J 4000 =. .... lOO t _ mOmxmd 11ncD_ 24100 i I I I

o_o4 o.so 0oO0 0.64 1.00 1 .SO 2.00 2.S0 $.00 S.S0 4.00

__) Time 8rim _ (me) (a) _)

44;0 1200

700

400 I ooo _r_.

f_ soo J 500 j I00 ,, 400 $00

I I I i200 _ '°° // /Y I/f _ Gmad Tin4 Doll 1oo | I ! I I I

0.00 OoSO 1.00 I .SO 2.00 2.S0 3.00 S.50 4.@0 0.1_ O.H 1 .(HI l.Se _.llO 2.M :I.O0 3+S0 4.00

800 f \ ,oo f

,oo y_ isoo I.. J i 400 "'_ !.o// 1

_ 200

too ...... Ground Tul OmM

I I I

1.04 ! .80 2.00 2.50 3.00 S.SO 4.00

11me Imm _s_ _) (c) Figure 8 RLIOA-3.3A Start Transient Simulation Output for Ground Test Conditions

17 SO0 IN00 I I 14000

$00

i --,v:-_ (,.il_?mspbIgslsDm (/ 400 mnl_q)

j:lO0 I I I /jr J II,000 // i iI0ql+ 0

200 4000

_100 / __.Jl_

I l J

0 O.S 1 1.5 2 2.S 0.5 I 1.5 2 2.5 3

_k,i, _d_ (re, I) 'lnil m,,ml MBS fJ,O

Figure 9 RL10A-3-3A Start Transient Simulation Output for Atlas/Centaur Flight AC-72 (first burn)

4100 II0_

140_ SO0 -- " - ,-i...._ 4oo _"_'+"+ I I J J 300

200 I il- _ 100 l I

0.04} 0.SO 1.10 $ .SO 2.04 2.S0 3,.00 _1.$4} 4.N

arm ]dgl.S (lie)

Figure 10 RLI0A-3-3A Start Transient Slmulatio_ Output for Arias/Centaur Flight AC-74 (2rid burn)

18 Table 2 Comparisonof Measuredand Predicted RL10 Engine Time-to-Accelerate

Type of Run Simulation Time Measured Time diFFerence (sec from MES) (sec from MES) (msec)

Cm3m_ Test 2.09 2.26 170 (early)

Grm_ Test 1.80 1.90 100 (early)

C._md Test 1.51 1.43 80 (late)

Cam_ Test 1.72 1.70 20 (late)

Grmmd Test (Relight) 1.91 1.84 70 (late)

Caotmd Test (Religh0 2.00 2.08 80 (early)

Flight 1.98 1.90 80 (late)

Flight 1.95 1.67 280 (late) * _t O_ugho 2.33 2.56 230(early)

* Note : Although this run had inlet conditions similar to other flights, these engines started about 300 msec earlier relative to MF.S. This may indicate a difference in the engine other than inlet conditions (see section ofthisreporton uncertainty).

I000

1700 _ _

Z I

J 1500

1SO@ J J | 11oo r

O00 !

i 700 X

500

O.O00.SO 1.001JO2.002.SOS.005.f_4.00 Tim, ZrmmDam _m,)

Figure 11 Predicted Maximum Cooling Jacket Metal Temperature during Engine Start

19 S_ 104)00 "_1 I I I I 450 11HlO0

...... i_.md TIll OlI I

I_00 j,. ; 1..

SO S@OO

0.2 0.4 0.1 O.I 1 0.2 0.4 O.I 0.8 tSmtmmJkdnhsJ_Cea-ddmd(l_CO)(m) TS/e hm Mdn helm C_eJ..dr lisnd OiIK_O) Ime) (a) (b)

I00 7@0 I -.---._,-- (s.tn.) Lolml m J--_ --_-_-.- . " Jlkm _Qmmml TIIII Bill

1.°° j _ 300 k L.\\ i,oo 20O ,. _'_ ,% _,. \ _,,_.----_ _.qqW_ _ 0 0

0 I).2 0.4 0.1 O.I I 0.2 0.4 O.I O.l

"nine tn,m Mde _ C:w-e8 dpd (JdU:O) (me) _m _ u-L. icqh, (:_ dpd IMICO) ira) ((:) (d) Figure 12 RLIOA-3-3A Shutdown Simulation for Ground Test Conditions

2O Form Approved REPORT DOCUMENTATION PAGE OMBNo.0704-0188

Public reporling burden for this collection of information is estimaled to average 1 hour per response, including the time for reviewing instructions, searching existing data sources. gathering and maintaining the data noecled, and cornple!ing a.nd reviewing the co.lleclion of inJormation:. Send _ts r,egar.ding th_bu_enestimateorany ot_her.,1Sa_jeff_h_s cofleclion o| informalion, including suggestions for reducing this burden, to Wash=ngton Headquarters :_ervees, Uweczorale lot mTormal_on ul_eratlunr* cmu net-_,_:,. --:_ ,Jw _,-, Davis Highway. Suite 1204, Arfington. VA 22202.4302. and to the Office of Management and Budget. Paperwork Reduction Prolect (0704-0188). Washinglon. DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED June 1995 Final Contractor Report

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS A Transient Model of the RL10A-3-3A Rocket Engine

WU-242-70--04 6. AUTHOR(S) C-NAS3-27186 Michael P. Binder

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER NYMA, Inc. Engineering Services Division E-9702 2001 Aerospace Parkway Brook Park, Ohio 44142

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER

National Aeronautics and Space Adminisu'ation NASA CR-195478 Lewis Research Center AIAA-95-2968 Cleveland, Ohio 44135-3191

11. SUPPLEMENTARYNOTES Prepared for the 31st Joint Propulsion Conference and Exhibit, cosponsored by AIAA, ASME, SAE, ASEE, San Diego, California, July 10-12, 1995. Project manager, Joseph A. Hemminger, Space Propulsion Technology Division, NASA Lewis Research Center, organization code 5320, (216) 433-7563. 12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified - Unlimited Subject Category 20

This publication is available from the NASA Center for Aerospace Information, (301) 621---0390. 13. ABSTRACT (Maximum200 words)

RL10A-3-JA rocket engines have served as the main propulsion system for Centaur upper stage vehicles since the early 1980's. This hydrogen/oxygen expander cycle engine continues to play a major role in the American launch industry. The Space Propulsion Technology Division at the NASA Lewis Research Center has created a computer model of the RL10 engine, based on detailed component analyses and available test data. This R10 engine model can predict the performance of the engine over a wide range of operating conditions. The model may also be used to predict the effects of any proposed design changes and anticipated failure scenarios. In this paper, the results of the component analyses are discussed. Simulation results from the new system model are compared with engine test and flight data, including the start and shut-down transient characteristics.

14. SUBJECTTERMS 15. NUMBER OF PAGES 22 RL10; RL10A3-3A; Centaur; Transient system model; ROCETS Program 16. PRICE CODE A03

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified Standard Form 298 (Rev. 2-89) NSN 7540-01-280-5500 Presc_'iloecl by ANSI Std. Z39-18 298-102

Cl_ • I_ -L ill Q. O CD

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