Historical Evolution of the Space Shuttle Primary and Vernier Reaction Control Rocket Engine Designs

2006 ◽  
Author(s):  
Carl Stechman ◽  
Charity Lawson
Keyword(s):  
Space Shuttle ◽  
Rocket Engine ◽  
Historical Evolution

scholarly journals Transmittance and Radiance Computations for Rocket Engine Plume Environments

2003 ◽  
Author(s):  
Gopal D. Tejwani
Keyword(s):  
Molecular Species ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Line Of Sight ◽  
Layer By Layer ◽  
Atmospheric Conditions ◽  
Combustion Gas ◽  
Exhaust Plume ◽  
Temperature And Pressure ◽  
High Degree

Rocket engine exhaust plume is generally thermal in character arising from changes in the internal energy of constituent molecules. Radiation from the plume is attenuated in its passage through the atmosphere. In the visible and the infrared region of the spectrum for clear-sky conditions, this is caused mainly through absorption by atmospheric molecular species. The most important combustion-product molecules giving rise to emission in the IR are water vapor, carbon dioxide, and carbon monoxide. In addition, the high temperature plume reacting with the surrounding atmosphere may produce nitrogen oxides, in the boundary layer, all of which are strongly emitting molecules. Important absorbing species in the atmosphere in the engine plume environment are H2O, CO, CO2, CH4, N2O, NO, and NO2. Under normal atmospheric conditions, the concentrations of O3, SO2, and NH3 are too small to produce any significant absorption. Essentially the problem comprises of the propagation of radiation from a hot gas source through a long cool absorbing atmosphere thus combining aspects of atmospheric and combustion gas methods. Since many of the same molecular species are responsible for both emission and absorption, the high degree of line position correlation between the emission and absorption spectra precludes the decoupling of the optical path into isolated emitter and absorber regions and multiplying the source band radiance by the absorber band transmittance in order to arrive at the transmitted radiance spectrum. Also, very strong thermal gradients may be encountered. All this suggests that a layer-by-layer computation is called for. The pathlength through the plume and the atmosphere is assumed to go through a certain number of layers, each of which is considered to have all molecular species in local thermodynamic equilibrium at constant temperature and pressure within the layer. Radiative transfer problems can be visualized as a set of parallel layers orthogonal to the line of sight, each with an input radiance from the previous layer and an output radiance to the subsequent layer. The MODTRAN (MODerate resolution TRANsmission) code is ideally suited for layer-by-layer absorption/emission calculations for atmospheric molecular species. We have utilized MODTRAN 4.0 computer code, implemented on a Power Mac G3, for the radiance and transmittance computations. The MODTRAN code has been adapted for the engine plume radiance computations. If the plume composition and flowfield parameters such as the temperature and pressure values are known along the line of sight by means of the experimental measurements or (more likely) CFD simulations, one can compute the radiance from any plume with high degree of accuracy at any desired point in space. Emission and absorption characteristics of several atmospheric and combustion species have been studied and presented in this paper with reference to the rocket engine plume environments at the Stennis Space Center. In general transmittance losses can not be neglected for any pathlength of 2 m or more. We have also studied the effect of clouds, rain, and fog on the plume radiance/transmittance. The transmittance losses are severe if any of these occur along the line of sight. Preliminary results for the radiance from the exhaust plume of the space shuttle main engine are shown and discussed.

Life extending controller design for reusable rocket engines

2001 ◽  
Vol 105 (1048) ◽  
pp. 315-322 ◽  
Author(s):  
A. Ray ◽  
M. S. Holmes ◽  
C. F. Lorenzo
Keyword(s):  
Controller Design ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Design Procedure ◽  
Fuel Mixture ◽  
Significant Loss ◽  
Chamber Pressure ◽  
Control Synthesis ◽  
Rocket Engines ◽  
Mixture Ratio

Abstract The goal of life extending control (LEC) is to enhance structural durability of complex mechanical systems, such as aircraft, spacecraft, and energy conversion devices, without incurring any significant loss of performance. This paper presents a concept of robust life-extending controller design for reusable rocket engines, similar to the Space Shuttle Main Engine (SSME), via damage mitigation in both fuel and oxidiser turbines while achieving the required performance for transient responses of the main combustion chamber pressure and the oxidant/fuel mixture ratio. The design procedure makes use of a combination of linear robust control synthesis and nonlinear optimisation techniques. Results of simulation experiments on the model of a reusable rocket engine are presented to this effect.

Cold Flow Testing of the Space Shuttle Main Engine Alternate Turbopump Development High Pressure Fuel Turbine Model

1992 ◽  
Author(s):  
Stephen W. Gaddis ◽  
Susan T. Hudson ◽  
P. Dean Johnson
Keyword(s):  
High Pressure ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Performance Model ◽  
Test Article ◽  
Turbine Performance ◽  
Reynolds Number Effects ◽  
Liquid Rocket ◽  
Turbine Design ◽  
Main Engine

The National Aeronautics and Space Administration’s (NASA’s) Marshall Space Flight Center (MSFC) has established a “cold” airflow turbine test program to experimentally determine the performance of liquid rocket engine turbopump drive turbines. Testing of the space shuttle main engine (SSME) alternate turbopump development (ATD) fuel turbine was conducted for “back-to-back” comparisons with the baseline SSME fuel turbine results obtained in the first quarter of 1991. Turbine performance, Reynolds number effects, and turbine diagnostics, such as stage reactions and exit swirl angles, were investigated at the turbine design point and at off-design conditions. The test data showed that the ATD fuel turbine test article was approximately 1.4 percent higher in efficiency and flowed 5.3 percent more than the baseline fuel turbine test article. This paper describes the method and results used to validate the ATD fuel turbine aerodynamic design. The results are being used to determine the ATD high pressure fuel turbopump (HPFTP) turbine performance over its operating range, anchor the SSME ATD steady-state performance model, and validate various prediction and design analyses.

Radial Inlet and Exit Design for a 10 MWe sCO2 Axial Turbine

2019 ◽  
Author(s):  
Stefan D. Cich ◽  
J. Jeffrey Moore ◽  
Michael Marshall ◽  
Kevin Hoopes ◽  
Jason Mortzheim ◽  
...  
Keyword(s):  
Power Density ◽  
High Efficiency ◽  
Environmental Changes ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Thermal Mass ◽  
Steam Turbines ◽  
Axial Turbine ◽  
Mechanical Constraints ◽  
Liquid Rocket

Abstract An enabling technology for a successful deployment of the sCO2 closed-loop recompression Brayton cycle is the development of a high temperature turbine not currently available in the marketplace. This turbine was developed under DOE funding for the STEP Pilot Plant development and represents a second generation design of the Sunshot turbine (Moore, et al., 2018). The lower thermal mass and increased power density of the sCO2 cycle, as compared to steam-based systems, enables the development of compact, high-efficiency power blocks that can respond quickly to transient environmental changes and frequent start-up/shut-down operations. The power density of the turbine is significantly greater than traditional steam turbines and is rivaled only by liquid rocket engine turbo pumps, such as those used on the Space Shuttle Main Engines. One key area that presents a design challenge is the radial inlet and exit collector to the axial turbine. Due to the high power density and overall small size of the machine, the available space for this inlet, collectors and transition regions is limited. This paper will take a detailed look at the space constraints and also the balance of aero performance and mechanical constraints in designing optimal flow paths that will improve the overall efficiency of the cycle.

scholarly journals Numerical Analyses of a Rocket Engine Turbine and Comparison With Air Test Data

1992 ◽  
Author(s):  
Ken Tran ◽  
Daniel C. Chan ◽  
Susan T. Hudson ◽  
Stephen W. Gaddis
Keyword(s):  
High Pressure ◽  
Test Data ◽  
Space Flight ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Second Stage ◽  
Overall Performance ◽  
Main Engine ◽  
Level Of Analysis ◽  
Nasa Marshall Space

Cold air test data on the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) turbine were recently collected at NASA Marshall Space Flight Center (MSFC). The turbine is a two-stage reaction machine, which was designed in the early 1970s (Fig. 1a). Overall performance data, static pressures on the first- and second-stage nozzles, and static pressures along the gas path at the hub and tip were gathered and are compared in this paper with various (1-D, quasi 3-D, and 3-D viscous) analysis procedures. The results of each level of analysis is compared to test data to demonstrate the range of applicability for each step in the design process of a turbine.

Twice-Running-Speed Response Due to Elliptical Bearing Clearances

2003 ◽  
Vol 125 (1) ◽  
pp. 64-67 ◽  
Author(s):  
Dara W. Childs
Keyword(s):  
Natural Frequency ◽  
Technology Development ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Development Program ◽  
Advanced Technology ◽  
Running Speed ◽  
Dead Band ◽  
Speed Range ◽  
Speed Response

Yamamoto [1] examined a vertical Jeffcott rotor model with bearing clearances and showed that a small bearing “dead-band” clearance could have a dramatic influence on synchronous rotor response. Recent test results for a turbopump in a liquid-rocket-engine development program showed a twice-running-speed response (2E) on the order of 8 gs over a wide speed range, while synchronous response levels were only at an 0.5 g level. Childs [2] predicted a sharp 2E response for the Advanced Technology Development, High-Pressure Fuel Turbopump (ATD-HPFTP) of the Space Shuttle Main Engine due to bearing clearances, when the running speed was nearly one half of a housing-mode natural frequency, i.e., when the 2E frequency was close to the housing mode natural frequency. However, his model did not predict significant 2E response over a broad running-speed range. Ellipticity of the bearing dead-band clearances was suggested as a possible cause for the observed 2E phenomenon. An extension of Yamamoto’s analysis [1] is presented including bearing ellipticity to examine that proposed explanation. The analysis results show that clearance ellipticity will produce 2E response over a considerable running-speed range during which the bearing clearance is engaged; however, the predicted 2E-response amplitude corresponding to an ellipticity of 0.1 were about 10% of the synchronous levels, and the projected 2E acceleration levels were about 40% of synchronous. The predicted 2E response includes a resonance peak (that can be sharp) at speeds slightly above 25% of the rotor critical speed. The perturbation-analysis results provide an explanation for persistent, lower 2E-response levels observed in many turbopumps, but do not explain the high levels observed with this turbopump.

Twice-Running-Speed Response due to Elliptical Bearing Clearances

2001 ◽  
Author(s):  
Dara W. Childs
Keyword(s):  
Natural Frequency ◽  
Technology Development ◽  
Space Shuttle ◽  
Rocket Engine ◽  
Development Program ◽  
Advanced Technology ◽  
Running Speed ◽  
Dead Band ◽  
Speed Range ◽  
Speed Response

Abstract Yamamoto [1] examined a vertical Jeffcott rotor model with bearing clearances and showed that a small bearing “dead-band” clearance could have a dramatic influence on synchronous rotor response. Recent test results for a turbopump in a liquid-rocket-engine development program showed a twice-running-speed response (2E) on the order of 8 gs over a wide speed range, while synchronous response levels were only at an 0.5 g level. Childs [2] predicted a sharp 2E response for the Advanced Technology Development, High-Pressure Fuel Turbopump of the Space Shuttle Main Engine due to bearing clearances, when the running speed was nearly one half of a housing-mode natural frequency, i.e., when the 2E frequency was close to the housing mode natural frequency. However, his model did not predict significant 2E response over a broad running-speed range. Ellipticity of the bearing dead-band clearances was suggested as a possible cause for the observed 2E phenomenon. An extension of Yamamoto’s analysis [1] is presented including bearing ellipticity to examine that proposed explanation. The analysis results show that clearance ellipticity will produce 2E response over a considerable running-speed range during which the bearing clearance is engaged; however, the predicted 2E-response amplitude corresponding to an ellipticity of 0.1 were about 10% of the synchronous levels, and the projected 2E acceleration levels were about 40% of synchronous. The predicted 2E response includes a resonance peak (that can be sharp) at speeds slightly above 25% of the rotor critical speed. The perturbation-analysis results provide an explanation for persistent, lower 2E-response levels observed in many turbopumps, but do not explain the high levels observed with this turbopump.

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