Design of injectors and ablative thrust chambers for a flox-propane rocket engine with a 1.2-inch throat diameter
Design of injectors and ablative thrust chambers for FLOX propane rocket engine with 1.2 inch throat diameter
Pavli, A. J.↗
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Advanced cooling techniques for high-pressure, hydrocarbon-fueled rocket engines
The regenerative cooling limits (maximum chamber pressure) are defined for oxygen/hydrocarbon (Methane, Propane, and RP-1) rocket engines over a thrust range of 20,000 to 600,000 lbf for a reusable life of 250 missions. Chamber pressure limits are first defined without a hot-gas wall carbon layer (unenhanced designs). Cooling enhancement chamber pressure limits are then established for seven thermal barriers (carbon layer, ceramic coating, graphite liner, film cooling, zoned combustion, transpiration cooling, and a combination of two of the above). The maximum regenerative-cooled chamber pressure is attained with the oxygen/methane propellant combination.
Cook, R. T.↗
Plume flowfield analysis of the shuttle primary Reaction Control System (RCS) rocket engine
A solution was generated for the physical properties of the Shuttle RCS 4000 N (900 lb) rocket engine exhaust plume flowfield. The modeled exhaust gas consists of the five most abundant molecular species, H2, N2, H2O, CO, and CO2. The solution is for a bare RCS engine firing into a vacuum; the only additional hardware surface in the flowfield is a cylinder (=engine mount) which coincides with the nozzle lip outer corner at X = 0, extends to the flowfield outer boundary at X = -137 m and is coaxial with the negative symmetry axis. Continuum gas dynamic methods and the Direct Simulation Monte Carlo (DSMC) method were combined in an iterative procedure to produce a selfconsistent solution. Continuum methods were used in the RCS nozzle and in the plume as far as the P = 0.03 breakdown contour; the DSMC method was used downstream of this continuum flow boundary. The DSMC flowfield extends beyond 100 m from the nozzle exit and thus the solution includes the farfield flow properties, but substantial information is developed on lip flow dynamics and thus results are also presented for the flow properties in the vicinity of the nozzle lip.
Hueser, J. E.↗
Modification of the PARC Navier-Stokes Code to predict rocket engine nozzle performance
The PARC2D Navier-Stokes Code was modified to compute the performance parameters for rocket engine nozzles. The perfect gas code was applied to the SSME engine nozzle for inviscid, laminar and turbulent flow. Inviscid computations compare well with Rocketdyne computations. Performance degradation due to the boundary layers is very reasonable. Application of the code to nontraditional nozzle geometries and to low Reynolds nozzles is demonstrated. Modification of the code for equilibrium H2/O2 chemistry is described. Thermodynamic and equilibrium constants are determined from statistical mechanics and the transport properties from exact kinetic theory, using collison integrals determined from appropriate intermolecular potentials. The equilibrium code was used to compute the SSME flowfield. Modifications of the flowfield due to the change of composition are described.
Collins, Frank G.↗
Development of a combustor analytical design methodology for liquid rocket engines
The development of a user friendly computerized methodology for the design and analysis of liquid propellant rocket engine combustion chambers is described. An overview of the methodology, consisting of a computer program containing an appropriate modular assembly of existing industry wide performance and combustion stability models, is presented. These models are linked with an interactive front end processor enabling the user to define the performance and stability traits of an existing design (point analysis) or to create the essential design features of a combustor to meet specific performance goals and combustion stability (point design). Plans for demonstration and verification of this methodology are also presented. These plans include the creation of combustor designs using the methodology, together with predictions of the performance and combustion stability for each design. A verification test program of 26 hot fire tests with up to four designs created using this methodology is described. This testing is planned using LOX/RP-1 propellants with a thrust level of approx. 220,000 N (50,000 lbf).
Pieper, Jerry L.↗
Ceramic Matrix Composite Turbine Disk for Rocket Engines
NASA has recently completed testing of a ceramic matrix composite (CMC), integrally bladed disk (blisk) for rocket engine turbopumps. The turbopump's main function is to bring propellants from the tank to the combustion chamber at optimal pressures, temperatures, and flow rates. Advantages realized by using CMC blisks are increases in safety by increasing temperature margins and decreasing costs by increasing turbopump performance. A multidisciplinary team, involving materials, design, structural analysis, nondestructive inspection government, academia, and industry experts, was formed to accomplish the 4.5 year effort. This article will review some of the background and accomplishments of the CMC Blisk Program relative to the benefits of this technology.
Effinger, Mike↗
Parallel Unsteady Turbopump Simulations for Liquid Rocket Engines
This paper reports the progress being made towards complete turbo-pump simulation capability for liquid rocket engines. Space Shuttle Main Engine (SSME) turbo-pump impeller is used as a test case for the performance evaluation of the MPI and hybrid MPI/Open-MP versions of the INS3D code. Then, a computational model of a turbo-pump has been developed for the shuttle upgrade program. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Time-accuracy of the scheme has been evaluated by using simple test cases. Unsteady computations for SSME turbo-pump, which contains 136 zones with 35 Million grid points, are currently underway on Origin 2000 systems at NASA Ames Research Center. Results from time-accurate simulations with moving boundary capability, and the performance of the parallel versions of the code will be presented in the final paper.
Kiris, Cetin C.↗
Analytical Redundancy Using Kalman Filters for Rocket Engine Sensor Validation
The use of sensor redundancy is crucial in aerospace systems to maintain safe, reliable operation. While hardware redundancy is more common in application, analytical redundancy can provide a viable alternative in systems where the installation of multiple redundant sensors is not viable. To this end, the use of Kalman filters to analytically validate sensor measurements within rocket engines was explored. First, a dynamic model of the RS 25 engine, a derivative of the Space Shuttle Main Engine (SSME), was reduced to a subset of relations, focused around the main combustion chamber pressure. These relations were used within the Kalman filter algorithm to generate an estimate of sensor measurements to be compared with true measurements for data validation purposes. By using a bank of Kalman filters, the residuals between the estimated and true measurements were used to detect and isolate sensor faults. Through fault simulations, the sensor validation performance of this Kalman filter bank design was compared to a hardware redundancy check. Sensor bias and drift faults of various magnitudes were injected into nominal RS 25 engine test data. Results for both approaches show comparable fault detection with most bias faults found nearly instantaneously by both algorithms. Drift fault detection results show certain cases where one algorithm is faster than the other. The key advantage of the Kalman filter algorithm is shown in fault isolation performance where it can isolate faults between two redundant sensors while the hardware redundancy comparisons cannot.
sensors↗
Near-Field Acoustical Characterization of Clustered Rocket Engines
This paper presents an approach for the prediction and characterization of the near-field acoustic levels from closely-spaced clustered rocket engines. The calculations are based on the method proposed by Eldred, wherein the flowfield from the clustered rockets is divided into two zones. Zone 1 contains the isolated nozzles which produce noise independently, and extends up to a distance where the individual flows completely mix to form an equivalent single nozzle flow. Zone 2 is occupied by the single mixed stream starting from the station where the jets merge. The acoustic fields from the two zones are computed separately on the basis of the NASA-SP method of Eldred developed for a single equivalent nozzle. A summation of the spectra for the two zones yields the total effective sound pressure level for the clustered engines. Under certain conditions of nozzle spacing and flow parameters, the combined sound pressure level spectrum for the clustered nozzles displays a double peak. Test cases are presented here to demonstrate the importance of hydrodynamic interactions responsible for the double peak in the sound spectrum in the case of clustered rocket nozzles, and the role of ground reflections in the case of non-interfering jets. A graphics interface (Rocket Acoustic Prediction Tool) has been developed to take into account the effects of clustered nozzles and ground reflections.
Kandula, Max↗
Elimination of rocket engine asymmetric loads during tests at sea level
Secondary injection concept eliminates asymmetric loads and may increase thrust rocket engine loads during sea level tests. The concept uses either a tubular manifold with evenly spaced injection ports or secondary fluid injected at the turbine exhaust inlet to the thrust chamber.
Johnson, J. R.↗
The open-cycle gas-core nuclear rocket engine - Some engineering considerations.
A preliminary design study of a conceptual 6000-MW open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 44,200 lb and a specific impulse of 4400 sec. The nuclear fuel is uranium-235 and the propellant is hydrogen. Critical fuel mass was calculated for several reactor configurations. Major components of the reactor (reflector, pressure vessel) and the waste heat rejection system were considered conceptually and were sized.
Taylor, M. F.↗
Digital controller for high pressure rocket engine.
Description of a general approach for the design of an adaptive digital control system for liquid bipropellant rocket engines. The technique employs linearized transfer functions derived from perturbations of an engine simulation. The linear models serve as a basis on which to develop candidate closed-loop control laws quickly and economically.
Thompson, Z.↗
FLOX/MMH rocket engine assemblies
A 6-month program was conducted to fabricate, acceptance test, and deliver to JPL two FLOX/MMH rocket engine assemblies. Acceptance testing consisted of injector cold flows and 15-second-duration hot firings with the engine assemblies. An additional objective of the program was to determine the feasibility of utilizing FLOX/MMH thrust chamber components with the F2/N2H4 propellant combination. Feasibility studies consisted of analysis, cold flow, and comparative F2/N2H4-FLOX/MMH calorimeter hot fire testing, followed by a 500-second duration F2/N2H4 demonstration test.
Source record↗
Combustion dynamics in liquid rocket engines
A chemical analysis of the emission and absorption spectra in the combustion chamber of a nitrogen tetroxide/aerozine-50 rocket engine was conducted. Measurements were made under conditions of preignition, ignition, and post combustion operating periods. The cause of severe ignition overpressures sporadically observed during the vacuum startup of the Apollo reaction control system engine was investigated. The extent to which residual propellants or condensed intermediate reaction products remain after the engine has been operated in a pulse mode duty cycle was determined.
Mclain, W. H.↗
RETSCP-A computer program for analysis of rocket engine thermal strains with cyclic plasticity
Finite element program employs three-dimensional isoparametric element for analysis of rocket engine thermal strains with cyclic plasticity.
Miller, R. W.↗
Technology status of a liquid fluorine-hydrazine rocket engine for a planetary spacecraft propulsion system
This paper discusses the current status of a fluorine-hydrazine rocket engine development program. Incorporation of a thin rhenium inner liner successfully eliminated corrosion of the carbon/carbon composite thrust chamber wall experienced during a previous test program. The results of hot-fire tests utilizing reworked and new injectors which provide increased fuel film cooling showed that thrust chamber head-end temperatures could be maintained at an acceptable level. As expected, the accompanying specific impulse performance loss requires optimizing the amount of film cooling to minimize the loss. The efforts to refine the rhenium liner vapor deposition process culminated in a carbon/carbon composite thrust chamber total test duration of 1008 seconds. Tasks presently in process include: (1) fabrication of two carbon/carbon composite thrust chambers incorporating 60:1 expansion ratio nozzles; (2) injector tests to optimize performance and cooling; (3) additional refinements to the rhenium lining process; and (4) fabrication and test of a freestanding rhenium thrust chamber.
Appel, M. A.↗
Quantifying Instability Sources in Liquid Rocket Engines
Computational fluid dynamics methodology to predict the effects of combusting flows on acoustic pressure oscillations in liquid rocket engines (LREs) is under development. 'Me intent of the investigation is to develop the causal physics of combustion driven acoustic resonances in LREs. The crux of the analysis is the accurate simulation of pressure/density/sound speed in a combustor which when used by the FDNS-RFV CFD code will produce realistic flow phenomena. An analysis of a gas generator considered for the Fastrac engine will be used as a test validation case.
Farmer, Richard C.↗
A Design Tool for Liquid Rocket Engine Injectors
A practical design tool which emphasizes the analysis of flowfields near the injector face of liquid rocket engines has been developed and used to simulate preliminary configurations of NASA's Fastrac and vortex engines. This computational design tool is sufficiently detailed to predict the interactive effects of injector element impingement angles and points and the momenta of the individual orifice flows and the combusting flow which results. In order to simulate a significant number of individual orifices, a homogeneous computational fluid dynamics model was developed. To describe sub- and supercritical liquid and vapor flows, the model utilized thermal and caloric equations of state which were valid over a wide range of pressures and temperatures. The model was constructed such that the local quality of the flow was determined directly. Since both the Fastrac and vortex engines utilize RP-1/LOX propellants, a simplified hydrocarbon combustion model was devised in order to accomplish three-dimensional, multiphase flow simulations. Such a model does not identify drops or their distribution, but it does allow the recirculating flow along the injector face and into the acoustic cavity and the film coolant flow to be accurately predicted.
Farmer, R.↗