Mechanical design of a Curtis turbine for the oxidizer turbopump of the M-1 engine
Turbine mechanical design for M-1 engine oxidizer turbopump
M-1 ROCKET ENGINE↗
SEARCH · Engineering Papers
Results for “Rocket Engine Design”
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
STS-45 Space Shuttle mission report
The STS-45 Space Shuttle Program Mission Report contains a summary of the vehicle subsystem operations during the forty-sixth flight of the Space Shuttle Program and the eleventh flight of the Orbiter Vehicle Atlantis (OV-104). In addition to the Atlantis vehicle, the flight vehicle consisted of the following: an External Tank (ET) designated as ET-44 (LWT-37); three Space Shuttle main engines (SSME's), which were serial numbers 2024, 2012, and 2028 in positions 1, 2, and 3, respectively; and two Solid Rocket Boosters (SRB's) designated as BI-049. The lightweight redesigned Solid Rocket Motors (RSRM's) installed in each of the SRB's were designated as 360L021A for the left SRM and 360W021B for the right SRM. The primary objective of this mission was to successfully perform the planned operations of the Atmospheric Laboratory for Applications and Science-1 (ATLAS-1) and the Shuttle Solar Backscatter Ultraviolet Instrument (SSBUV) payloads. The secondary objectives were to successfully perform all operations necessary to support the requirements of the following: the Space Tissue Loss-01 (STL-01) experiment; the Radiation Monitoring Equipment-3 (RME-3) experiment; the Visual Function Tester-2 (VFT-2) experiment; the Cloud Logic to Optimize use of Defense System (CLOUDS-1A) experiment; the Shuttle Amateur Radio Experiment 2 (SAREX-2) Configuration B; the Investigation into Polymer Membranes Processing experiment; and the Get-Away Special (GAS) payload G-229. The Ultraviolet Plume Instrument (UVPI) was a payload of opportunity that required no special maneuvers. In addition to the primary and secondary objectives, the crew was tasked to perform as many as 10 Development Test Objectives (DTO'S) and 14 Detailed Supplementary Objectives (DSO's).
Fricke, Robert W.↗
STS-42 Space Shuttle mission report
The STS-42 Space Shuttle Program Mission Report contains a summary of the vehicle subsystem operations during the forty-fifth flight of the Space Shuttle Program and the fourteenth flight of the Orbiter vehicle Discovery (OV-103). In addition to the Discovery vehicle, the flight vehicle consisted of the following: an External Tank (ET) designated as ET-52 (LWT-45); three Space Shuttle main engines (SSME's), which were serial numbers 2026, 2022, and 2027 in positions 1, 2, and 3, respectively; and two Solid Rocket Boosters (SRB's) designated as BI-048. The lightweight redesigned Solid Rocket Motors (RSRM's) installed in each one of the SRB's were designated as 360L020A for the left SRM and 360Q020B for the right SRM. The primary objective of the STS-42 mission was to complete the objectives of the first International Microgravity Laboratory (IML-1). Secondary objectives were to perform all operations necessary to support the requirements of the following: Gelation of Sols: Applied Microgravity Research (GOSAMR); Student Experiment 81-09 (Convection in Zero Gravity); Student Experiment 83-02 (Capillary Rise of Liquid Through Granular Porous Media); the Investigation into Polymer Membrane Processing (IPMP); the Radiation Monitoring Equipment-3 (RME-3); and Get-Away Special (GAS) payloads carried on the GAS Beam Assembly.
Fricke, Robert W.↗
Pros, Cons, and Alternatives to Weight Based Cost Estimating
Many cost estimating tools use weight as a major parameter in projecting the cost. This is often combined with modifying factors such as complexity, technical maturity of design, environment of operation, etc. to increase the fidelity of the estimate. For a set of conceptual designs, all meeting the same requirements, increased weight can be a major driver in increased cost. However, once a design is fixed, increased weight generally decreases cost, while decreased weight generally increases cost - and the relationship is not linear. Alternative approaches to estimating cost without using weight (except perhaps for materials costs) have been attempted to try to produce a tool usable throughout the design process - from concept studies through development. This paper will address the pros and cons of using weight based models for cost estimating, using liquid rocket engines as the example. It will then examine approaches that minimize the imp~ct of weight based cost estimating. The Rocket Engine- Cost Model (RECM) is an attribute based model developed internally by Pratt & Whitney Rocketdyne for NASA. RECM will be presented primarily to show a successful method to use design and programmatic parameters instead of weight to estimate both design and development costs and production costs. An operations model developed by KSC, the Launch and Landing Effects Ground Operations model (LLEGO), will also be discussed.
Joyner, Claude R.↗
Pressure measurement
Practices are outlined for the design, installation, checkout, calibration, and operation of a pressure measuring system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) pressure measurement system elemental uncertainties; (2) short- and long-term pressure measurement system uncertainty; (3) shunt calibration of pressure transducers; (4) special considerations for vacuum measurement; and (5) methods of determining the dynamic characteristics of pressure transducers. Design guidelines are provided for the critical components of each portion of the system to provide a pressure measurement system which meets the performance criteria specified.
Source record↗
Free-flight Performance of a Rocket-boosted, Air-launched 16-inch-diameter Ram-jet Engine at Mach Numbers up to 2.20
The investigation of air-launched ram-jet engines has been extended to include a study of models with a nominal design free-stream Mach number of 2.40. These models require auxiliary thrust in order to attain a flight speed at which the ram jet becomes self-accelerating. A rocket-boosting technique for providing this auxiliary thrust is described and time histories of two rocket-boosted ram-jet flights are presented. In one flight, the model attained a maximum Mach number of 2.20 before a fuel system failure resulted in the destruction of the engine. Performance data for this model are presented in terms of thrust and drag coefficients, diffuser pressure recovery, mass-flow ratio, combustion efficiency, specific fuel consumption, and over-all engine efficiency.
AIR INLETS - NOSE, ANNULAR↗
On the X-34 FASTRAC-Memorandums of Misunderstanding
Engineers at MSFC designed, developed, and tested propulsion systems that helped launch Saturn I, IB, and V boosters for the Apollo missions. After the Apollo program, Marshall was responsible for the design and development of the propulsion elements for the Shuttle launch vehicle, including the solid rocket boosters, external tank and main engines. Each of these systems offered new propulsion technological challenges that pushed engineers and administrators beyond Saturn. The technical challenges presented by the development of each of these propulsion systems helped to establish and sustain a culture of engineering conservatism and was often accompanied by a deep level of penetration into contractors that worked on these systems.
Hawkins, Lakiesha V.↗
The Role of the Strutjet Engine in New Global and Space Markets
The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling propulsion concept for single stage to orbit applications. Recent design considerations indicate that this systems also provides benefits supportive of other commercial non-space applications. This paper describes the technical progress of the Strutjet since 1997 together with a rationale why Rocket Based Combined Cycle Engines in general, and the Strutjet in particular, lend themselves uniquely to systems having the ability to expand current space and open new global 'rapid delivery' markets. During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design maturity has been continuously improved and desired features, like simple variable geometry and low drag flowpath resulting in high performance, have been verified. In addition, data is now available which allows the designer, who is challenged to maximize system operability and economic feasibility, to choose between hydrogen or hydrocarbon fuels for a variety of application. The ability exists now to apply this propulsion system to various vehicles with a multitude of missions. In this paper, storable hydrocarbon and gaseous hydrogen Strutjet RBCC test data as accomplished to date and as planned for the future is presented, and the degree of required technology maturity achieved so far is assessed. Two vehicles, using cryogenic propane fuel Strutjet engines, and specifically designed for rapid point-to-point cargo delivery between Pacific rim locations are introduced, discussed, and compared.
Siebenhaar, A.↗
The Role of the Strutjet Engine in New Global and Space Markets
The Strutjet, discussed in previous IAF papers, was originally introduced as an enabling propulsion concept for single stage to orbit applications. Recent design considerations indicate that this systems also provides benefits supportive of other commercial non-space applications. This paper describes the technical progress of the Strutjet since 1997 together with a rationale why Rocket Based Combined Cycle Engines in general, and the Strutjet in particular, lend themselves uniquely to systems having the ability to expand current space and open new global "rapid delivery" markets. During this decade, Strutjet technology has been evaluated in over 1000 tests. Its design maturity has been continuously improved and desired features, like simple variable geometry and low drag flowpath resulting in high performance, are verified. In addition, data is now available which allows the designer, who is challenged to maximize system operability and economic feasibility, to choose between hydrogen or hydrocarbon fuels for these application. The ability exists now to apply this propulsion system to various vehicles with a multitude of missions. In this paper, the previously presented earth-to-orbit hydrogen powered vehicle is up3ated and another vehicle, specifically designed for rapid point-to-point delivery, is introduced and discussed. High payoff propulsion technologies required for these vehicles are identified and laid out in a roadmap spanning over the next decade.
Siebenhaar, A.↗
Reactor moderator, pressure vessel, and heat rejection system of an open-cycle gas core nuclear rocket concept
A preliminary design study of a conceptual 6000-megawatt open-cycle gas-core nuclear rocket engine system was made. The engine has a thrust of 196,600 newtons (44,200 lb) and a specific impulse of 4400 seconds. 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 waste heat rejection system) were considered conceptually and were sized.
Taylor, M. F.↗
Advanced oxygen-hydrocarbon rocket engine study
Preliminary identification and evaluation of promising liquid oxygen/ hydrocarbon (LO2/HC) rocket engine cycles is reported. A consistent and reliable data base for vehicle optimization and design studies, to demonstrate the significance of propulsion system improvements, and to select the critical technology areas necessary to realize such advances is presented.
Obrien, C. J.↗
Engineering and programming manual: Two-dimensional kinetic reference computer program (TDK)
The Two Dimensional Kinetics (TDK) computer program is a primary tool in applying the JANNAF liquid rocket thrust chamber performance prediction methodology. The development of a methodology that includes all aspects of rocket engine performance from analytical calculation to test measurements, that is physically accurate and consistent, and that serves as an industry and government reference is presented. Recent interest in rocket engines that operate at high expansion ratio, such as most Orbit Transfer Vehicle (OTV) engine designs, has required an extension of the analytical methods used by the TDK computer program. Thus, the version of TDK that is described in this manual is in many respects different from the 1973 version of the program. This new material reflects the new capabilities of the TDK computer program, the most important of which are described.
Nickerson, G. R.↗
Oxidizer heat exchanger component testing
As part of the RL10 Rocket Engine Product Improvement Program, Oxidizer Heat Exchanger (OHE) stages 1, 2, and 3 were designed and fabricated during late 1983 and early 1984. The purpose of the OHE is to provide gaseous oxygen to the propellant injector for stable engine operation at tank head idle and pumped idle operating modes. This report summarizes the OHE stages 1 and 3 rig testing, and includes the separation of the stage 1-and-2 assembly and the remanifolding of stage 1. The OHE performance analysis and analytical model modifications for both stages are also presented. The flow tests were accomplished during the time period from 9 October 1984 to 12 November 1984.
Kmiec, T.↗
Symposium (International) on Combustion, 20th, University of Michigan, Ann Arbor, MI, August 12-17, 1984, Proceedings
The present conference on combustion phenomena considers topics in automotive engine combustion, turbulent reacting flows, the modeling of practical combustion systems, reaction kinetics, combustion-generated particulates, combustion diagnostics, coal combustion process characteristics, fire-related phenomena, explosion/detonation phenomena, spray combustion, ignition/extinction, laminar flames, pollutant formation processes, practical combustor devices, and rocket propellant combustion. Attention is given to the contributions of combustion science to piston engine design, modeling and measurement techniques for turbulent combustion, the specific effects of energy, collisions, and transport processes in combustion chemistry kinetics, the formation of large molecules, particulates and ions in premixed hydrocarbon flames, the application of laser diagnostics to combustion systems, spark ignition energies for dust-air mixtures, the controlling mechanisms of flow-assisted flame spread, the ignition and combustion of coal-water slurries, spontaneous ignition of methane, turbulent and accelerating dust flames, and the temperature sensitivity of double base propellants.
Source record↗
Finite element based electric motor design optimization
The purpose of this effort was to develop a finite element code for the analysis and design of permanent magnet electric motors. These motors would drive electromechanical actuators in advanced rocket engines. The actuators would control fuel valves and thrust vector control systems. Refurbishing the hydraulic systems of the Space Shuttle after each flight is costly and time consuming. Electromechanical actuators could replace hydraulics, improve system reliability, and reduce down time.
Campbell, C. Warren↗
Subsonic Glideback Rocket Demonstrator Flight Testing
For the past two years, Cal Poly's rocket program has been aggressively exploring the concept of remotely controlled, fixed wing, flyable rocket boosters. This program, embodied by a group of student engineers known as Cal Poly Space Systems, has successfully demonstrated the idea of a rocket design that incorporates a vertical launch pattern followed by a horizontal return flight and landing. Though the design is meant for supersonic flight, CPSS demonstrators are deployed at a subsonic speed. Many steps have been taken by the club that allowed the evolution of the StarBooster prototype to reach its current size: a ten-foot tall, one-foot diameter, composite material rocket. Progress is currently being made that involves multiple boosters along with a second stage, third rocket.
DeTurris, Dianne J.↗
Incompressible Navier-Stokes Computations for the Development of a Ventricular Assist Device
An incompressible flow analysis code, INS3D, has been applied to the development of a mechanical heart assist device. The solution method is based on the artificial compressibility approach and uses an implicit-upwind differencing scheme together with a Gauss-Seidel line relaxation method. The equations are solved in steadily rotating reference frames and the centrifugal and the Coriolis force terms are included as source terms. The resulting computational procedure is validated for liquid rocket engine analysis and applied subsequently to analyze a Ventricular Assist Device (VAD). A new design configuration is developed which includes an inducer upstream of the impeller main blades, and substantial improvement is observed in the performance of the VAD.
Kiris, Cetin↗
Updated Mars Mission Architectures Featuring Nuclear Thermal Propulsion
Nuclear thermal propulsion (NTP) can potentially enable routine human exploration of Mars and the solar system. By using nuclear fission instead of a chemical combustion process, and using hydrogen as the propellant, NTP systems promise rocket efficiencies roughly twice that of the best chemical rocket engines currently available. The most recent major Mars architecture study featuring NTP was the Design Reference Architecture 5.0 (DRA 5.0), performed in 2009. Currently, the predominant transportation options being considered are solar electric propulsion (SEP) and chemical propulsion; however, given NTP's capabilities, an updated architectural analysis is needed. This paper provides a top-level overview of several different architectures featuring updated NTP performance data. New architectures presented include a proposed update to the DRA 5.0 as well as an investigation of architectures based on the current Evolvable Mars Campaign, which is the focus of NASA's current analyses for the Journey to Mars. Architectures investigated leverage the latest information relating to NTP performance and design considerations and address new support elements not available at the time of DRA 5.0, most notably the Orion crew module and the Space Launch System (SLS). The paper provides a top level quantitative comparison of key performance metrics as well as a qualitative discussion of improvements and key challenges still to be addressed. Preliminary results indicate that the updated NTP architectures can significantly reduce the campaign mass and subsequently the costs for assembly and number of launches.
Rodriguez, Mitchell A.↗