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At least 1,135 records · Page 63
All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Metal Matrix Composites for Rocket Engine Applications

This document is from a presentation about the applications of Metal Matrix Composites (MMC) in rocket engines. Both NASA and the Air Force have goals which would reduce the costs and the weight of launching spacecraft. Charts show the engine weight distribution for both reuseable and expendable engine components. The presentation reviews the operating requirements for several components of the rocket engines. The next slide reviews the potential benefits of MMCs in general and in use as materials for Advanced Pressure Casting. The next slide reviews the drawbacks of MMCs. The reusable turbopump housing is selected to review for potential MMC application. The presentation reviews solutions for reusable turbopump materials, pointing out some of the issues. It also reviews the development of some of the materials.

McDonald, Kathleen R.↗
NASA NTRS2000

Development of Ceramic Matrix Composite Turbine Blisks for Rocket Engines

Ceramic matrix composite (CMC) integrally bladed turbine disks (blisks) are being considered for use in various advanced propulsion systems for space vehicles. The successful development of this technology can significantly impact National Aeronautics and Space Administration (NASA) space transportation missions, by enabling new efficient systems that can operate at higher temperatures, while reducing costs. Composite blisks comprised of carbon (C) fibers and a silicon carbide (SiC) ceramic matrix were designed, fabricated, characterized, and tested by a multidisciplinary team involving materials, design, structural analysis, turbomachinery, and nondestructive evaluation representatives from government, academia, and industry during a 4.5 year effort led by the NASA Marshall Space Flight Center (MSFC). The testing of several of these blisks, which were developed in the Simplex Turbopump CMC Blisk ]Program for use in rocket engine turbopumps, was recently completed. CMC blisks offer potential advantages in rocket engine turbopumps including increased safety resulting from increased operating temperature margins and greater pump reliability, and decreased costs resulting from improved turbopump performance. The progress that was achieved in that development effort is reviewed, and some of the technology that could be applied to other advanced space transportation propulsion systems is discussed.

Effinger, Mike↗
NASA NTRS1997

On the Development of a Magnetically Vectored Variable ISP Plasma Rocket

The development of a Magnetically Vectored Variable I(sub sp) Plasma Rocket at the Advanced Space Propulsion Laboratory (ASPL) is in progress at NASA's Johnson Space Center. The facility is using a small, 3.2 m tandem mirror device to study the application of RF heated magnetically contained plasmas for space propulsion. The central cell radius is 0.1 m and fields of 0.2 T and 2 T are possible in the central and end-cell mirror sections, respectively. A magnetoplasmadynamic (MPD) injector has just been acquired and will be used along with other methods of plasma refueling. A 1 MW magnet power supply upgrade is being developed with full implementation by the Spring of 1997. Two microwave systems for discharge initiation and plasma heating at 2.45 GHz and 14.0 GHz, respectively, are in operation. Additionally, RF systems with 200 kW and 1 MW of power are being modified and conditioned for operation. The concept provides electrode-less operation and variable thrush'specific impulse at constant power (200 -30 N /5000-30,000 seconds at 10 MW with a 60% efficiency). Optimization for speed or payload are possible with the same engine, giving the rocket great flexibility. Missions to Mars in 90 days are described, and missions to Pluto are under study.

Feliciano, Enectali Figueroa↗
NASA NTRS2000

FDNS CFD Code Benchmark for RBCC Ejector Mode Operation: Continuing Toward Dual Rocket Effects

Computational Fluid Dynamics (CFD) analysis results are compared with benchmark quality test data from the Propulsion Engineering Research Center's (PERC) Rocket Based Combined Cycle (RBCC) experiments to verify fluid dynamic code and application procedures. RBCC engine flowpath development will rely on CFD applications to capture the multi -dimensional fluid dynamic interactions and to quantify their effect on the RBCC system performance. Therefore, the accuracy of these CFD codes must be determined through detailed comparisons with test data. The PERC experiments build upon the well-known 1968 rocket-ejector experiments of Odegaard and Stroup by employing advanced optical and laser based diagnostics to evaluate mixing and secondary combustion. The Finite Difference Navier Stokes (FDNS) code [2] was used to model the fluid dynamics of the PERC RBCC ejector mode configuration. Analyses were performed for the Diffusion and Afterburning (DAB) test conditions at the 200-psia thruster operation point, Results with and without downstream fuel injection are presented.

West, Jeff↗
NASA NTRS2000

Evaluation of Vortex Chamber Concepts for Liquid Rocket Engine Applications

Rocket-based combined-cycle engines (RBBC) being considered at NASA for future generation launch vehicles feature clusters of small rocket thrusters as part of the engine components. Depending on specific RBBC concepts, these thrusters may be operated at various operating conditions including power level and/or propellant mixture ratio variations. To pursue technology developments for future launch vehicles, NASA/Marshall Space Flight Center (MSFC) is examining vortex chamber concepts for the subject cycle engine application. Past studies indicated that the vortex chamber schemes potentially have a number of advantages over conventional chamber methods. Due to the nature of the vortex flow, relatively cooler propellant streams tend to flow along the chamber wall. Hence, the thruster chamber can be operated without the need of any cooling techniques. This vortex flow also creates strong turbulence, which promotes the propellant mixing process. Consequently, the subject chamber concepts not only offer the system simplicity but they also would enhance the combustion performance. The test results showed that the chamber performance was markedly high even at a low chamber length-to- diameter ratio (L/D). This incentive can be translated to a convenience in the thrust chamber packaging.

Trinh, Huu Phuoc↗
NASA NTRS2001

Robust Low Cost Liquid Rocket Combustion Chamber by Advanced Vacuum Plasma Process

Next-generation, regeneratively cooled rocket engines will require materials that can withstand high temperatures while retaining high thermal conductivity. Fabrication techniques must be cost efficient so that engine components can be manufactured within the constraints of shrinking budgets. Three technologies have been combined to produce an advanced liquid rocket engine combustion chamber at NASA-Marshall Space Flight Center (MSFC) using relatively low-cost, vacuum-plasma-spray (VPS) techniques. Copper alloy NARloy-Z was replaced with a new high performance Cu-8Cr-4Nb alloy developed by NASA-Glenn Research Center (GRC), which possesses excellent high-temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability. Functional gradient technology, developed building composite cartridges for space furnaces was incorporated to add oxidation resistant and thermal barrier coatings as an integral part of the hot wall of the liner during the VPS process. NiCrAlY, utilized to produce durable protective coating for the space shuttle high pressure fuel turbopump (BPFTP) turbine blades, was used as the functional gradient material coating (FGM). The FGM not only serves as a protection from oxidation or blanching, the main cause of engine failure, but also serves as a thermal barrier because of its lower thermal conductivity, reducing the temperature of the combustion liner 200 F, from 1000 F to 800 F producing longer life. The objective of this program was to develop and demonstrate the technology to fabricate high-performance, robust, inexpensive combustion chambers for advanced propulsion systems (such as Lockheed-Martin's VentureStar and NASA's Reusable Launch Vehicle, RLV) using the low-cost VPS process. VPS formed combustion chamber test articles have been formed with the FGM hot wall built in and hot fire tested, demonstrating for the first time a coating that will remain intact through the hot firing test, and with no apparent wear. Material physical properties and the hot firing tests are reviewed.

Holmes, Richard↗
NASA NTRS2001

Rocket Launch-Induced Vibration and Ignition Overpressure Response

Rocket-induced vibration and ignition overpressure response environments are predicted in the low-frequency (5 to 200 hertz) range. The predictions are necessary to evaluate their impact on critical components, structures, and facilities in the immediate vicinity of the rocket launch pad.

Caimi, Raoul E.↗
NASA NTRS2001

A Low Cost GPS System for Real-Time Tracking of Sounding Rockets

In an effort to minimize the need for costly, complex, tracking radars, the German Space Operations Center has set up a research project for GPS based tracking of sounding rockets. As part of this project, a GPS receiver based on commercial technology for terrestrial applications has been modified to allow its use under the highly dynamical conditions of a sounding rocket flight. In addition, new antenna concepts are studied as an alternative to proven but costly wrap-around antennas.

Markgraf, M.↗
NASA NTRS2001

Acoustical and Flowfield Characterization of a Scaled Tabletop Rocket

An analysis of the acoustical and flowfield environment for the scaled 1-pound-force (lbf) thrust tabletop motor was performed. This tabletop motor from NASA Stennis Space Center Is composed of Plexiglas burning In gaseous oxygen with a graphite insert for the nozzle portion. The nozzle has a throat diameter of 0.2 inch and an exit diameter of 0.38 Inch. With a chamber pressure at 55 pounds per square Inch absolute (psia), a normal shock is formed immediately downstream of the nozzle exit plane as the combustion products exhaust into the ambient at atmospheric pressure. The jet characterization Is based on computational fluid dynamics (CFD) in conjunction with Kirchhoff surface integral formulation and compared with correlations developed for measured rocket noise and a pressure fluctuation scaling (PFS) method. Predictions and comparisons are made for the overall sound pressure levels (OASPL's) and spectral dependence of sound pressure level (SPL). The overall objective of this effort is to develop methods for scaling the acoustic and flowfield environment of rockets with a wide range of thrust (1 lbf to 1 million lbf).

Kandula, Max↗
NASA NTRS2001

Multiobjective Optimization of Rocket Engine Pumps Using Evolutionary Algorithm

A design optimization method for turbopumps of cryogenic rocket engines has been developed. Multiobjective Evolutionary Algorithm (MOEA) is used for multiobjective pump design optimizations. Performances of design candidates are evaluated by using the meanline pump flow modeling method based on the Euler turbine equation coupled with empirical correlations for rotor efficiency. To demonstrate the feasibility of the present approach, a single stage centrifugal pump design and multistage pump design optimizations are presented. In both cases, the present method obtains very reasonable Pareto-optimal solutions that include some designs outperforming the original design in total head while reducing input power by one percent. Detailed observation of the design results also reveals some important design criteria for turbopumps in cryogenic rocket engines. These results demonstrate the feasibility of the EA-based design optimization method in this field.

Oyama, Akira↗
NASA NTRS2001

Robust Strategy for Rocket Engine Health Monitoring

Monitoring the health of rocket engine systems is essentially a two-phase process. The acquisition phase involves sensing physical conditions at selected locations, converting physical inputs to electrical signals, conditioning the signals as appropriate to establish scale or filter interference, and recording results in a form that is easy to interpret. The inference phase involves analysis of results from the acquisition phase, comparison of analysis results to established health measures, and assessment of health indications. A variety of analytical tools may be employed in the inference phase of health monitoring. These tools can be separated into three broad categories: statistical, rule based, and model based. Statistical methods can provide excellent comparative measures of engine operating health. They require well-characterized data from an ensemble of "typical" engines, or "golden" data from a specific test assumed to define the operating norm in order to establish reliable comparative measures. Statistical methods are generally suitable for real-time health monitoring because they do not deal with the physical complexities of engine operation. The utility of statistical methods in rocket engine health monitoring is hindered by practical limits on the quantity and quality of available data. This is due to the difficulty and high cost of data acquisition, the limited number of available test engines, and the problem of simulating flight conditions in ground test facilities. In addition, statistical methods incur a penalty for disregarding flow complexity and are therefore limited in their ability to define performance shift causality. Rule based methods infer the health state of the engine system based on comparison of individual measurements or combinations of measurements with defined health norms or rules. This does not mean that rule based methods are necessarily simple. Although binary yes-no health assessment can sometimes be established by relatively simple rules, the causality assignment needed for refined health monitoring often requires an exceptionally complex rule base involving complicated logical maps. Structuring the rule system to be clear and unambiguous can be difficult, and the expert input required to maintain a large logic network and associated rule base can be prohibitive.

Santi, L. Michael↗
NASA NTRS2001

Rocket Launch-Induced Vibration and Ignition Overpressure Response

Rocket-induced vibration and ignition overpressure response environments are predicted in the low-frequency (5 to 200 hertz) range. The predictions are necessary to evaluate their impact on critical components, structures, and facilities in the immediate vicinity of the rocket launch pad.

Caimi, Raoul↗
NASA NTRS2001

Laser Ignition Technology for Bi-Propellant Rocket Engine Applications

This viewgraph presentation gives an overview of laser ignition technology for bipropellant rocket engines applications. The objectives of this project include: (1) the selection test chambers and flows; (2) definition of the laser ignition setup; (3) pulse format optimization; (4) fiber optic coupled laser ignition system analysis; and (5) chamber integration issues definition. The testing concludes that rocket combustion chamber laser ignition is imminent. Support technologies (multiplexing, window durability/cleaning, and fiber optic durability) are feasible.

Thomas, Matt↗
NASA NTRS2001

Internal Flow Simulation of Enhanced Performance Solid Rocket Booster for the Space Transportation System

An enhanced performance solid rocket booster concept for the space shuttle system has been proposed. The concept booster will have strong commonality with the existing, proven, reliable four-segment Space Shuttle Reusable Solid Rocket Motors (RSRM) with individual component design (nozzle, insulator, etc.) optimized for a five-segment configuration. Increased performance is desirable to further enhance safety/reliability and/or increase payload capability. Performance increase will be achieved by adding a fifth propellant segment to the current four-segment booster and opening the throat to accommodate the increased mass flow while maintaining current pressure levels. One development concept under consideration is the static test of a "standard" RSRM with a fifth propellant segment inserted and appropriate minimum motor modifications. Feasibility studies are being conducted to assess the potential for any significant departure in component performance/loading from the well-characterized RSRM. An area of concern is the aft motor (submerged nozzle inlet, aft dome, etc.) where the altered internal flow resulting from the performance enhancing features (25% increase in mass flow rate, higher Mach numbers, modified subsonic nozzle contour) may result in increased component erosion and char. To assess this issue and to define the minimum design changes required to successfully static test a fifth segment RSRM engineering test motor, internal flow studies have been initiated. Internal aero-thermal environments were quantified in terms of conventional convective heating and discrete phase alumina particle impact/concentration and accretion calculations via Computational Fluid Dynamics (CFD) simulation. Two sets of comparative CFD simulations of the RSRM and the five-segment (IBM) concept motor were conducted with CFD commercial code FLUENT. The first simulation involved a two-dimensional axi-symmetric model of the full motor, initial grain RSRM. The second set of analyses included three-dimensional models of the RSRM and FSM aft motors with four-degree vectored nozzles.

Ahmad, Rashid A.↗
NASA NTRS2002

Robust Low Cost Liquid Rocket Combustion Chamber by Advanced Vacuum Plasma Process

Next-generation, regeneratively cooled rocket engines require materials that can meet high temperatures while resisting the corrosive oxidation-reduction reaction of combustion known as blanching, the main cause of engine failure. A project was initiated at NASA-Marshal Space Flight Center (MSFC) to combine three existing technologies to build and demonstrate an advanced liquid rocket engine combustion chamber that would provide a 100 mission life. Technology developed in microgravity research to build cartridges for space furnaces was utilized to vacuum plasma spray (VPS) a functional gradient coating on the hot wall of the combustion liner as one continuous operation, eliminating any bondline between the coating and the liner. The coating was NiCrAlY, developed previously as durable protective coatings on space shuttle high pressure fuel turbopump (HPFTP) turbine blades. A thermal model showed that 0.03 in. NiCrAlY applied to the hot wall of the combustion liner would reduce the hot wall temperature 200 F, a 20% reduction, for longer life. Cu-8Cr-4Nb alloy, which was developed by NASA-Glenn Research Center (GRC), and which possesses excellent high temperature strength, creep resistance, and low cycle fatigue behavior combined with exceptional thermal stability, was utilized as the liner material in place of NARloy-Z. The Cu-8Cr-4Nb material exhibits better mechanical properties at 650 C (1200 F) than NARloy-Z does at 538 C (1000 F). VPS formed Cu-8Cr-4Nb combustion chamber liners with a protective NiCrAlY functional gradient coating have been hot fire tested, successfully demonstrating a durable coating for the first time. Hot fire tests along with tensile and low cycle fatigue properties of the VPS formed combustion chamber liners and witness panel specimens are discussed.

Holmes, Richard↗
NASA NTRS2002

Blood Pump Development Using Rocket Engine Flow Simulation Technology

This viewgraph presentation provides information on the transfer of rocket engine flow simulation technology to work involving the development of blood pumps. Details are offered regarding the design and requirements of mechanical heart assist devices, or VADs (ventricular assist device). There are various computational fluid dynamics issues involved in the visualization of flow in such devices, and these are highlighted and compared to those of rocket turbopumps.

Kiris, Cetin C.↗
NASA NTRS2002

Rocket Launch Trajectory Simulations Mechanism

The design and development of a Trajectory Simulation Mechanism (TSM) for the Launch Systems Testbed (LST) is outlined. In addition to being one-of-a-kind facility in the world, TSM serves as a platform to study the interaction of rocket launch-induced environments and subsequent dynamic effects on the equipment and structures in the close vicinity of the launch pad. For the first time, researchers and academicians alike will be able to perform tests in a laboratory environment and assess the impact of vibroacoustic behavior of structures in a moving rocket scenario on ground equipment, launch vehicle, and its valuable payload or spacecraft.

Margasahayam, Ravi↗
NASA NTRS1961

Performance Charts for Multistage Rocket Boosters

Charts relating the stage propellant fractions are given for two-and three-stage rockets launching payloads into nominal low-altitude circular orbits about the earth. A simple method is described for extending these data to higher orbit or escape missions. Various combinations of stages using RP - liquid-oxygen and hydrogen - liquid-oxygen propellants are considered. However, the results can be generalized with little error to any other propellant combination.Charts relating the stage propellant fractions are given for two-and three-stage rockets launching payloads into nominal low-altitude circular orbits about the earth. A simple method is described for extending these data to higher orbit or escape missions. Various combinations of stages using RP - liquid-oxygen and hydrogen - liquid-oxygen propellants are considered. However, the results can be generalized with little error to any other propellant combination.

MacKay, John S.↗