SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 18 - SEPTEMBER 09, 2005
20 SPACECRAFT PROPULSION AND POWER - PART III
Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxiliary power sources.
For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 15 Launch Vehicles and Launch Operations, and 44 Energy Production and Conversion.
20050207504 NASA Marshall Space Flight Center, Huntsville, AL, USA
EELV Booster Assist Options for CEV
McNeal, Curtis, Jr.; June 1, 2005; 14 pp.; In English; 41st AIAA/ASME/ASEE/SAE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: CASI; A03, Hardcopy
Medium lift EELVs may still play a role in manned space flight. To be considered for manned flight, medium lift EELVs must address the short comings in their current boost assist motors. Two options exist: redesign and requalify the solid rocket motors. Replace solid rocket motors (SRMs) with hybrid rocket motors. Hybrid rocket motors are an attractive alternative. They are safer than SRMs. The TRL’s Lockheed Martin Small Launch Vehicle booster development substantially lowers the development risk, cost risk, and the schedule risk for developing hybrid boost assist for EELVs. Hybrid boosters testability offsets SRMs higher inherent reliability.Hybrid booster development and recurring costs are lower than SRMs. Performance gains are readily achieved. Derived from text
Ducted Rocket Engines; Solid Propellant Rocket Engines; Launch Vehicles; Booster Rocket Engines
20050207508 NASA Marshall Space Flight Center, Huntsville, AL, USA
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Spacecraft Chemical Propulsion Systems at NASA’s Marshall Space Flight Center: Heritage and Capabilities
McRight, Patrick S.; Sheehy, Jeffrey A.; Blevins, John A.; [2005]; 7 pp.; In English; 41st AIAA/ASME/ASEE/SAE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: CASI; A02, Hardcopy
NASA Marshall Space Flight Center (MSFC) is well known for its contributions to large ascent propulsion systems such as the Saturn V and the Space Shuttle. This paper highlights a lesser known but equally rich side of MSFC - its heritage in spacecraft chemical propulsion systems and its current capabilities for in-space propulsion system development and chemical propulsion research. The historical narrative describes the efforts associated with developing upper-stage main propulsion systems such as the Saturn S-IVB as well as orbital maneuvering and reaction control systems such as the S-IVB auxiliary propulsion system, the Skylab thruster attitude control system, and many more recent activities such as Chandra, the Demonstration of Automated Rendezvous Technology, X-37, the X-38 de-orbit propulsion system, the Interim Control Module, the US Propulsion Module, and several technology development activities. Also discussed are MSFC chemical propulsion research capabilities, along with near- and long-term technology challenges to which MSFC research and system development competencies are relevant. Author
Spacecraft Propulsion; Propulsion System Configurations; Chemical Propulsion; Systems Engineering; Attitude Control; Flight Characteristics
20050207509 NASA Marshall Space Flight Center, Huntsville, AL, USA
Progress on the Plasmoid Thruster Experiment (PTX)
Martin, Adam; Eskridge, Richard; Fimognari, Peter; Koelfgen, Syri; [2005]; 1 pp.; In English; Joint Propulsion Conference, 11-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: Other Sources; Abstract Only
A plasmoid, also called a compact toroid, is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic field (Bp and Bt, respectively). An object with Bp/Bt much greater than 1 is called a Field Reverse Configuration (FRC); if Bp = Bt, it is called a Spheromak. The plasmoid thruster is a pulsed inductive device which operates by repetitively producing plasmoids that are accelerated and ejected at high velocity. As the process is inductive, this thruster avoids the problem of electrode erosion. Also, the magnetic structure of the plasmoid should suppress thermal and mass losses to the wall, and improve detachment of the plasma exhaust from the thruster. This concept should be capable of producing an Isp of 5,000 seconds and greater, with thrust densities of order 10(exp 5) N/sq m.
The plasmoid thruster consists chiefly of a conical theta-pinch coil. Propellant is introduced onto a bias magnetic field, produced by an auxiliary coil, and is then pre-ionized, freezing in the magnetic field. The theta-pinch coil is then energized producing a field aligned anti- parallel to the bias field. The reversed field reconnects with the bias field to form the plasmoid. The magnetic pressure of the reversed field accelerates the plasmoid out of the thruster.
A series of experiments have been conducted on the PTX device, which consisted of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank, which rang at a frequency of 500 kHz, and served all three functions required for formation: pre-ionization, bias field loading, and field reversal. Initial ionization was found to occur in an annular region at the exit plane of the coil, and was found to be reproducible with a variety of gases, including H2, D2, Ar, and an H2/N2 mixture (75% /25%). A fast gas valve for injecting propellant has been tested, as well as a ringing pre-ionization circuit (operating at 5 Mhz) to better control the plasmoid formation. Velocities as high as 20 km/s have been measured for a plasmoid propagating into a static back-fill of neutral gas. Magnetic field and density measurements indicate that the plasmoid retains its structure well into the down-stream exhaust region, and that it is capable of pushing a dense cloud of neutral gas in front of it. Further experiments will focus on the effects of coil geometry and bias-field strength on the performance of the device. A high-power solid-state switching system is being developed in order to replace the spark-gap switch. Author
Plasmas (Physics); Thrustors; Toroidal Plasmas; Magnetic Field Configurations
20050207548 NASA Marshall Space Flight Center, Huntsville, AL, USA
Heated Promoted Combustion-Initial Test Results
Engel, Carl D.; Herald, Stephen; Davis, S. Eddie; June 2005; 19 pp.; In English; National Space and Missile Materials Symposium, 27 Jun. - 1 Jul. 2005, Summerlin, NV, USA; Original contains color illustrations Contract(s)/Grant(s): NAS8-01050; No Copyright; Avail: CASI; A03, Hardcopy
The purpose of the STD 6001 test 17 is to determine the flammability of materials in GOX at ambient temperature and at use pressure. The purpose of the new Heated Promoted combustion test is to determine the flammability of material in GOX at use temperature and pressure. The objective is to present the new heated promoted combustion method and show initial data and trends for three representative metals. Derived from text
Combustion; Flammability; Performance Tests
20050207551 Boeing Co., Canoga Park, CA, USA
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Real-Time On-Board HMS/Inspection Capability for Propulsion and Power Systems
Barkhoudarian, Sarkis; [2005]; 6 pp.; In English; AIAA Conference, 10 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-01140; No Copyright; Avail: CASI; A02, Hardcopy
Presently, the evaluation of the health of space propulsion systems includes obtaining and analyzing limited flight data and extensive post flight performance, operational and inspection data.
This approach is not practical for deep-space missions due to longer operational times, lack of in-space inspection facility, absence of timely ground commands and very long repair intervals.
This paper identifies the on-board health- management/inspection needs of deep-space propulsion and thermodynamic power-conversion systems.
It also describes technologies that could provide on-board inspection and more comprehensive health management for more successful missions. Author
Systems Health Monitoring; Propulsion System Configurations; Real Time Operation; Inspection; Aerospace Systems
20050207570 NASA Marshall Space Flight Center, Huntsville, AL, USA
Experimental Investigation of Magnesium Powder Combustion with C02 for Mars Ascent Applications
Foote, John P.; Litchford, Ronald J.; [2005]; 1 pp.; In English; Joint Propulsion Conference, 11-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: Other Sources; Abstract Only
Combustion of metals with CO2 has been identified as a possible propellant for Mars ascent applications. CO2 could be condensed from the Martian atmosphere, reducing the amount of propellant that must be transported from Earth. An attractive feature of this approach compared to other in situ propellant concepts is that no chemical processing on Mars is required. Magnesium has been identified as the most promising metal for this application because it ignites and burns easily in CO2. Preliminary systems studies indicate a 2 to 1 delivered mass advantage for Mg ascent propulsion using in situ C02, as compared to a conventional storable propellant system. The Propulsion Research Center at MSFC is undertaking an experimental investigation of magnesium powder combustion with CO2 in order to provide fundamental data on the combustion performance of Mg powder + CO2 mixtures needed to assess the feasibility of developing a practical Mg powder + CO2 rocket engine. Initial combustion experiments will be carried out in a small scale atmospheric pressure dump combustor. Effects of varying the Mg particle size, firing rate and O/F ratio on combustion stability and efficiency will be investigated. The combustion process will be characterized by optical flame measurements and extraction of combustion product samples. The experimental facility is currently being prepared and combustion experiments will begin during the first quarter of 2005. The final paper will describe the test facility and initial experimental results. Author
Magnesium; Powder (Particles); Combustion Products; Carbon Dioxide
20050207571 NASA Marshall Space Flight Center, Huntsville, AL, USA
Microinstabilities in the Gasdynamic Mirror Propulsion System
Emrich, William; [2005]; 1 pp.; In English; Joint Propulsion Conference, 11-13 Jul. 2005, Tucson, AZ, USA; No Copyright; Avail: Other Sources; Abstract Only
The gasdynamic mirror has been proposed as a concept which could form the basis of a highly efficient fusion rocket engine. Gasdynamic mirrors differ from most other mirror type plasma confinement schemes in that they have much larger aspect ratios and operate at somewhat higher plasma densities. There are several types of instabilities which are known to plague mirror type confinement schemes. These instabilities fall into two general classes. One class of instability is the Magnetohydrodynamic or MHD instability which induces gross distortions in the plasma geometry. The other class of instability is the ‘loss cone’ microinstability which leads to general plasma turbulence. The ‘loss cone’ microinstability is caused by velocity space asymmetries resulting from the loss of plasma having constituent particle velocities within the angle of the magnetic mirror ‘loss cone.’ These instabilities generally manifest themselves in high temperature, moderately dense plasmas. The present study indicates that a GDM configured as a rocket engine might operate in a plasma regime where microinstabilities could potentially be significant. Author
Stability; Magnetohydrodynamic Stability; Magnetic Mirrors; Plasma Turbulence; Propulsion
20050207583 Alabama Univ., Huntsville, AL, USA
Investigation of Compressibility Effect for Aeropropulsive Shear Flows
Balasubramanyam, M. S.; Chen, C. P.; June 16, 2005; 7 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-14 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Contract(s)/Grant(s): NCC8-200 Report No.(s): AlAA Paper 2005-3712; No Copyright; Avail: CASI; A02, Hardcopy
Rocket Based Combined Cycle (RBCC) engines operate within a wide range of Mach numbers and altitudes. Fundamental fluid dynamic mechanisms involve complex choking, mass entrainment, stream mixing and wall interactions. The Propulsion Research Center at the University of Alabama in Huntsville is involved in an on- going experimental and numerical modeling study of non-axisymmetric ejector-based combined cycle propulsion systems. This paper attempts to address the modeling issues related to mixing, shear layer/wall interaction in a supersonic Strutjet/ejector flow field. Reynolds Averaged Navier-Stokes (RANS) solutions incorporating turbulence models are sought and compared to experimental measurements to characterize detailed flow dynamics. The effect of compressibility on fluids mixing and wall interactions were investigated using an existing CFD methodology. The compressibility correction to conventional incompressible twoequation models is found to be necessary for the supersonic mixing aspect of the ejector flows based on 2-D simulation results. 3-D strut-base flows involving flow separations were also investigated. Author
Compressibility Effects; Propulsion System Performance; Base Flow; Supersonic Flow; Mathematical Models; Computational Fluid Dynamics
20050209890 NASA Marshall Space Flight Center, Huntsville, AL, USA
REIMR: A Process for Utilizing Propulsion-Oriented ‘Lessons-Learned’ to Mitigate Development Risk
Ballard, Richard O.; Brown, Kendall K.; July 1, 2005; 9 pp.; In English; 41st Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-01108 Report No.(s): AIAA Paper 2005-4522; No Copyright; Avail: CASI; A02, Hardcopy
This paper is a summary overview of a study conducted a t the NASA Marshall Space Flight Center (MSFC) during the initial phases of the Space Launch Initiative (SLI) program to evaluate a large number of technical problems associated with the design, development, test, evaluation and operation of several major liquid propellant rocket engine systems (i.e., SSME, Fastrac, J-2, F-1). The results of this study was the identification of the ‘Fundamental Root Causes’ that enabled the technical problems to manifest, and practices that can be implemented to prevent them from recurring in future engine development efforts. This paper will discus the Fundamental Root Causes, cite some examples of how the technical problems arose from them, and provide a discussion of how they can be mitigated or avoided. Author
Liquid Propellant Rocket Engines; Engine Design; Space Shuttle Main Engine; Spacecraft Launching; J-2 Engine
20050209904 NASA Marshall Space Flight Center, Huntsville, AL, USA
Status of the Combustion Devices Injector Technology Program at the NASA MSFC
Jones, Gregg; Protz, Christopher; Trinh, Huu; Tucker, Kevin; Nesman, Tomas; Hulka, James; July 22, 2005; 20 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA; Original contains color illustrations Report No.(s): AIAA Paper 2005-4530; No Copyright; Avail: CASI; A03, Hardcopy
To support the NASA Space Exploration Mission, an in-house program called Combustion Devices Injector Technology (CDIT) is being conducted at the NASA Marshall Space Flight Center (MSFC) for the fiscal year 2005. CDIT is focused on developing combustor technology and analysis tools to improve reliability and durability of upper-stage and in-space liquid propellant rocket engines. The three areas of focus include injector/chamber thermal compatibility, ignition, and combustion stability. In the compatibility and ignition areas, small-scale single- and multi-element hardware experiments will be conducted to demonstrate advanced technological concepts as well as to provide experimental data for validation of computational analysis tools. In addition, advanced analysis tools will be developed to eventually include 3-dimensional and multi- element effects and improve capability and validity to analyze heat transfer and ignition in large, multi-element injectors. Derived from text
Combustion Chambers; Combustion Stability; Injectors; Heat Transfer
20050209928 Spectra Research Systems, Inc., Huntsville, AL, USA
Fabrication and Deployment Testing of Solar Sail Quadrants for a 20-Meter Solar Sail Ground Test System Demonstration
Laue, Greg; Case, David; Moore, Jim; June 03, 2005; 16 pp.; In English; 41st AIAA/ASME Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NAS8-27705; No Copyright; Avail: CASI; A03, Hardcopy
A 20-meter Scalable Square Solar Sail (S(sup 4)) System was produced and successfully completed functional vacuum testing in NASA Glenn’s Space Power Facility at Plum Brook Station Ohio in May 2005. The S(sup 4) system was designed and developed by ATK Space Systems, and the design and production of the Solar Sails for this system was carried out by SRS Technologies.
The S(sup 4) system consists of a central structure with four deployable carbon fiber masts that support four triangular sails. SRS has developed an effective and efficient design for triangular sail quadrants that are supported at three points and provide a flat reflective surface with a high fill factor. This sail design is robust enough for deployments in a one atmosphere, one gravity environment and incorporates several advanced features including adhesiveless seaming of membrane strips, compliant edge borders to allow for film membrane cord strain mismatch without causing wrinkling and low mass (3% of total sail mass) ripstop.
This paper will outline some of the sail design and fabrication processes and the mature production, packaging and deployment processes that have been developed. This paper will also detail the successful ambient and vacuum testing of the sails and the ATK spacecraft structure. Based on recent experience and testing, SRS is confidant that high Technology Readiness Level (TRL) 5-6 solar sails in the 40-120-meter size range with areal density in the 4-5 grams per square meters (sail minus structure) range can be produced with existing technology. Additional film production research will lead to further reductions in film thickness to less than 1 micron enabling production of sails with areal densities as low as 2.0 grams per square meters using the current design, resulting in a system areal densities as low as 5.3 grams per square meters (sail and structure). These areal densities are low enough to allow nearly all of the Solar Sail missions that have been proposed by the scientific community.
The fundamental technologies required to produce these systems has been demonstrated on the 20-meter S(sup 4) sails that have recently completed ground testing demonstrating a mature and technology suitable for incorporation into future flight validation and future mission. Solar Sails can support NASA’s Vision for Space Exploration by allowing communication satellite orbits that can maintain continuous communication with the polar regions of the Moon and Mars and to support solar weather monitoring to provide early warning of solar flares and storms that could threaten the safety of astronauts and other spacecraft. Author
Deployment; Fabrication; Ground Tests; Solar Sails; Aerospace Systems
20050209955 Sirius Group, Huntsville, AL, USA
Design Rules and Scaling for Solar Sails
Zeiders, Glenn W.; April 25, 2005; 11 pp.; In English; 41st AIAA Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NASA Order H-35186-D; No Copyright; Avail: CASI; A03, Hardcopy
Useful design rules and simple scaling models have been developed for solar sails. Chief among the conclusions are: 1. Sail distortions contribute to the thrust and moments primarily though the mean squared value of their derivatives (slopes), and the sail behaves like a flat sheet if the value is small. The RMS slope is therefore an important figure of merit, and sail distortion effects on the spacecraft can generally be disregarded if the RMS slope is less than about 10% or so. 2. The characteristic slope of the sail distortion varies inversely with the tension in the sail, and it is the tension that produces the principle loading on the support booms. The tension is not arbitrary, but rather is the value needed to maintain the allowable RMS slope. That corresponds to a halyard force about equal to three times the normal force on the supported sail area. 3. Both the AEC/SRS and L Garde concepts appear to be structurally capable of supporting sail sizes up to a kilometer or more with 1AU solar flux, but select transverse dimensions must be changed to do so. Operational issues such as fabrication, handling, storage and deployment will be the limiting factors. Author
Solar Sails; Spacecraft Design; Fabrication
20050209994 Naval Air Warfare Center, Patuxent River, MD, USA
Propulsion, Chapter 24
Thomas, Lawrence A.; Introduction to Flight Test Engineering, Volume 14; July 2005, pp. 24-1 - 24-14; In English; See also 20050209967; Copyright; Avail: CASI; A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document
This chapter provides a basic guide to the flight testing of propulsion system operability and compatibility (O&C). For the purposes of establishing a frame of reference, O&C refers to the ability of the aircrew to establish and maintain the desired level of propulsion system net propulsive force throughout the operating envelope of the aircraft. For the purposes of definition, net propulsive force is used to refer to the vector resultant of all throttle dependant forces acting upon the aircraft. By limiting the discussion in this Section to flight testing of the O&C of the propulsion system, it should not be interpreted to mean that these are the only factors that need to be considered when conducting flight test to evaluate an aircraft propulsion system. Propulsion system structural interfaces, pneumatic interfaces, mechanical interfaces, hydraulic interfaces, thermodynamic interfaces and electrical interfaces must all be evaluated prior to or concurrently with the O&C test program in order to ensure a safe and effective flight test program of the aircraft and propulsion system. Author
Flight Tests; Compatibility; Propulsion System Performance
20050210001 NASA Marshall Space Flight Center, Huntsville, AL, USA
Launch Vehicle Propulsion Design with Multiple Selection Criteria
Shelton, Joey D.; Frederick, Robert A.; Wilhite, Alan W.; [2005]; 28 pp.; In English; 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 10-13 Jul. 2005, Tucson, AZ, USA; Copyright; Avail: CASI; A03, Hardcopy
The approach and techniques described herein define an optimization and evaluation approach for a liquid hydrogen/liquid oxygen single-stage-to-orbit system. The method uses Monte Carlo simulations, genetic algorithm solvers, a propulsion thermo-chemical code, power series regression curves for historical data, and statistical models in order to optimize a vehicle system. The system, including parameters for engine chamber pressure, area ratio, and oxidizer/fuel ratio, was modeled and optimized to determine the best design for seven separate design weight and cost cases by varying design and technology parameters. Significant model results show that a 53% increase in Design, Development, Test and Evaluation cost results in a 67% reduction in Gross LiftoffWeight. Other key findings show the sensitivity of propulsion parameters, technology factors, and cost factors and how these parameters differ when cost and weight are optimized separately. Each of the three key propulsion parameters; chamber pressure, area ratio, and oxidizer/fuel ratio, are optimized in the seven design cases and results are plotted to show impacts to engine mass and overall vehicle mass. Author
Launch Vehicles; Liquid Hydrogen; Liquid Oxygen; Spacecraft Propulsion; Design Analysis; Pressure Ratio
20050210047 CU Aerospace, LLC, Champaign, IL, USA
UltraSail - Ultra-Lightweight Solar Sail Concept
Burton, Rodney L.; Coverstone, Victoria L.; Hargens-Rysanek, Jennifer; Ertmer, Kevin M.; Botter, Thierry; Benavides, Gabriel; Woo, Byoungsam; Carroll, David L.; Gierow, Paul A.; Farmer, Greg, et al.; April 27, 2005; 14 pp.; In English; 41st Joint Propulsion Conference, 10-13 Jul. 2005, Tucson, AZ, USA Contract(s)/Grant(s): NNM04AB18C; Copyright; Avail: CASI; A03, Hardcopy
UltraSail is a next-generation high-risk, high-payoff sail system for the launch, deployment, stabilization and control of very large (sq km class) solar sails enabling high payload mass fractions for high (Delta)V. Ultrasail is an innovative, non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation-flying micro-satellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 sq km, sail subsystem area densities approaching 1 g/sq m, and thrust levels many times those of ion thrusters used for comparable deep space missions. Ultrasail can achieve outer planetary rendezvous, a deep space capability now reserved for high-mass nuclear and chemical systems. One of the primary innovations is the near-elimination of sail supporting structures by attaching each blade tip to a formation-flying micro-satellite which deploys the sail, and then articulates the sail to provide attitude control, including spin stabilization and precession of the spin axis. These tip micro-satellites are controlled by 3-axis micro-thruster propulsion and an on-board metrology system. It is shown that an optimum spin rate exists which maximizes payload mass. Author
Solar Sails; Launching; Deployment; Spin Stabilization; Payloads; Formation Flying; Propulsion System Configurations; Attitude Control; Spacecraft Propulsion
Source: NASA.
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