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Mach 10 nozzle and 15 cu m (6' diameter x 15' long) dump chamber would provide a useful hypersonic facility for instruction and research.

GRA

pressure, gas stagnation temperature, tunnel wall structural temperature and flow Mach number are precisely controlled by the new controller in a stable manner. The tunnel control hardware, software, and the flow chart to assist in calibration of the sensors, actuators, and the controller real time functions are described. The software installation details are also presented. The report serves as the maintenance and trouble shooting manual for the 0.3-m TCT controller.

Author

12 ASTRONAUTICS (GENERAL)

For extraterrestrial exploration see 91 Lunar and Planetary Exploration.

N89-26870# Aktiebolaget Rollab, Stockholm (Sweden).
ON INTERNAL BENDING-BEAM STRAIN-GAGE WIND TUNNEL
BALANCES
Knut Fristedt Jan. 1989 99 p
(RR-070; ETN-89-94940) Avail: NTIS HC A05/MF A01

The relative load-carrying capacity of three types of internal balances was compared using information in the literature. The bending-beam balance is superior, when D is less than 35 mm. The capacity of task-balances is equal at larger diameters and the two-shell balance is a competitor, when D is greater than 50 mm. A strength computation model of the bending-beam balance of the Rollab type based on simple engineering theories is presented. The strength of available types of high alloy steels is discussed as well as the expected relations between mechanical stresses and electrical signals. It is shown during the calculation of the 30 mm balance, that the angle deflection of the sting at maximum load is very large, and that it is very beneficial to have access to a sting material with higher modulus of elasticity than steel. The influence of the numerical value of the diameter on the load carrying capacity at constant stress of the current type of balance is discussed, and task and two-shell balances are compared. A mathematical model of 2nd degree, which combines the force system loading the balance with the balance signals is described. This model can be used during calibration and wind-tunnel testing.

ESA

N89-26873*# National Aeronautics and Space Administration, Washington, DC. THE BEST WE CAN BE 1989 28 p (NASA-TM-101781; NAS 1.15:101781) Avail: NTIS HC A03/MF A01; SOD HC $2.00 as 033-000-01037-1 CSCL 22A

A brief history of the U.S. manned space program is presented covering the early Mercury and Gemini missions, the Apollo lunar missions, the Skylab project, and the Space Shuttle missions. The role of the Space Station in future manned missions is also discussed. Descriptions of government centers and major contractors supporting the Space Shuttle program are given.

M.G.

N89-26874# Deutsche Forschungsgemeinschaft, Bonn (Germany, F.R.). SPACE RESEARCH IN THE FEDERAL REPUBLIC OF GERMANY (FRG), 1986 AND 1987 1988 56 p (ETN-89-94427) Avail: NTIS HC A04/MF A01

Space astronomy and solar physics; solar system; solar wind, magnetospheres, and energetic particules; atmospheres and ionospheres; Earth observations; materials science; life sciences; and rocket launches are discussed.

ESA

13 ASTRODYNAMICS

Includes powered and free-flight trajectories; and orbital and launching dynamics.

8 p

N89-26871# Ballistic Research Labs., Aberdeen Proving Ground,
MD.
A MULTIDRIVER SHOCK TUBE MODEL OF A LARGE BLAST
SIMULATOR Final Report
Edmund J. Gion May 1989 41 p
(AD-A208324; BRL-MR-3757) Avail: NTIS HC A03/MF A01
CSCL 20/4

The construction of the BRL Multidriver Shock Tube (MD-ST) Model is documented. The facility is a 1:22 model of the Large Blast Simulator at the Centre d'Etude de Gramat, France, except for a lengthened driven tube to permit observation of the full waveform development without interference from the open-end, reflected rarefaction wave (since the Rarefaction Wave Eliminator was not modeled). It is designed with a good safety factor to withstand driver pressures to 24,000 kPa (3,500 psi). Initial tests to greater than 22,100 kPa (3,200 psi) have been performed, and results are compared to other available data. Additionally, the double diaphragm technique, described, was used to attain the highest shock pressures.

GRA

N89-26875# Instituto de Pesquisas Espaciais, Itajuba (Brazil).
FORMULATION OF NUMERICAL ORBIT PROPAGATION AND
TIME ERROR CORRECTION
Sebastiao Cicero Pinheiro Gomes and Helio Koitikuga Dec. 1988

Presented at the 39th Congress of the International Astronautical Federation, Bangalore, India, 8-15 Oct. 1988 (INPE-4774-PRE/1443) Avail: NTIS HC A02/MF A01

Several forms of numerical orbit propagation are reviewed. Basically, they consist of formulating in different fashions the differential equations which represent the dynamics of orbital motion. Such formulations (time transformation, stabilization, regularization) are developed within a fictitious time frame which varies nonlinearly with the physical time. Thus, the physical time becomes a generalized coordinate of the system which should be obtained numerically and, as a consequence, with a certain inevitable numerical error. The nature of this time error is analyzed and a simple correction is proposed here, together with computer simulation results.

Author

N89-26872# Naval Postgraduate School, Monterey, CA.
AN INVESTIGATION INTO THE USE OF AN EXISTING SHOCK
TUBE AS A DRIVER FOR A HYPERSONIC SHOCK TUNNEL
M.S. Thesis
Michael H. Sherman Mar. 1989 152 p
(AD-A208483) Avail: NTIS HC A08/MF A01 CSCL 01/1

This thesis describes experiments carried out using an existing tube alone, and with the tube connected to two-dimensional wedge nozzle. The range of maximum duration of steady reflected pressure from 3.5 to 5 milliseconds was achieved through tailored operation for incident shock strengths of 3.4 and 2.0, using pure Helium and a 70 percent Helium/30 percent Nitrogen mixture as the driver gas respectfully. Spark and continuous light shadowgraph techniques were attempted using an optical window at the Mach 4.3 location. Results demonstrated that the short duration flow phenomena in a shock tunnel can be recorded successfully using existing equipment. Calculations showed that the addition of a

14 GROUND SUPPORT SYSTEMS AND

FACILITIES (SPACE)

Includes launch complexes, research and production facilities; ground support equipment, e.g., mobile transporters; and simulators.

For related information see also 09 Research and Support Facilities (Air).

No abstracts in this category.

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17 SPACE COMM., SPACECRAFT

COMM., COMMAND & TRACKING

Includes telemetry; space communications networks; astronavigation and guidance; and radio blackout.

For related information see also 04 Aircraft Communications and Navigation and 32 Communications and Radar.

N89-26876*# National Aeronautics and Space Administration.
Lewis Research Center, Cleveland, OH.
NASA'S CHEMICAL TRANSFER PROPULSION PROGRAM FOR
PATHFINDER
Ned P. Hannum, Frank D. Berkopec, and Robert L. Zurawski
1989 18 p Presented at the 25th Joint Propulsion Conference,
Monterey, CA, 10-12 Jul. 1989; cosponsored by the AIAA, ASME,
SAE, and ASEE
(NASA-TM-102298; E-4976; NAS 1.15:102298; AIAA-89-2298)
Avail: NTIS HC A03/MF A01 CSCL 22B

Pathfinder is a research and technology project, with specific deliverables, initiated by the National Aeronautics and Space Administration (NASA) which will strengthen the technology base of the United States civil space program in preparation for future space exploration missions. Pathfinder begins in Fiscal Year 1989, and is to advance a collection of critical technologies for these missions and ensure technology readiness for future national decisions regarding exploration of the solar system. The four major thrusts of Pathfinder are: surface exploration, in-space operations, humans-in-space, and space transfer. The space transfer thrust will provide the critical technologies needed for transportation to, and return from, the Moon, Mars, and other planets in the solar system, as well as for reliable and cost-effective Earth-orbit operations. A key element of this thrust is the Chemical Transfer Propulsion program which will provide the propulsion technology for high performance, liquid oxygen/liquid hydrogen expander cycle engines which may be operated and maintained in space. Described here are the program overview including the goals and objectives, management, technical plan, and technology transfer for the Chemical Transfer Propulsion element of Pathfinder. Author

N89-26879*# Multipoint Communications Corp., Sunnyvale, CA. PROGRAMMABLE RATE MODEM UTILIZING DIGITAL SIGNAL PROCESSING TECHNIQUES Final Report George K. Bunya and Robert L. Wallace Jul. 1989 110 p (Contract NAS3-25336) (NASA-CR-185124; NAS 1.26:185124) Avail: NTIS HC A06/MF A01 CSCL 17B

The engineering development study to follow was written to address the need for a Programmable Rate Digital Satellite Modem capable of supporting both burst and continuous transmission modes with either binary phase shift keying (BPSK) or quadrature phase shift keying (QPSK) modulation. The preferred implementation technique is an all digital one which utilizes as much digital signal processing (DSP) as possible. Here design tradeoffs in each portion of the modulator and demodulator subsystem are outlined, and viable circuit approaches which are easily repeatable, have low implementation losses and have low production costs are identified. The research involved for this study was divided into nine technical papers, each addressing a significant region of concern in a variable rate modem design. Trivial portions and basic support logic designs surrounding the nine major modem blocks were omitted. In brief, the nine topic areas were: (1) Transmit Data Filtering; (2) Transmit Clock Generation; (3) Carrier Synthesizer; (4) Receive AGC; (5) Receive Data Filtering; (6) RF Oscillator Phase Noise; (7) Receive Carrier Selectivity; (8) Carrier Recovery; and (9) Timing Recovery.

Author

16 SPACE TRANSPORTATION

Includes passenger and cargo space transportation, e.g., shuttle operations; and space rescue techniques.

For related information see also 03 Air Transportation and Safety and 18 Spacecraft Design, Testing and Performance. For space suits see 54 Man/System Technology and Life Support

N89-26877*# National Aeronautics and Space Administration.
Lewis Research Center, Cleveland, OH.
SATELLITE RELOCATION BY TETHER DEPLOYMENT
Geoffrey A. Landis and Frank J. Hrach Apr. 1989 16 p
(NASA-TM-101992; E-4696; NAS 1.15:101992) Avail: NTIS HC
A03/MF A01 CSCL 22B

Several new uses of satellite tethers are discussed, including: (1) using tether extension to reposition a satellite in orbit without fuel expenditure by extending a mass on the end of a tether; (2) using a tether for energy storage to power the satellite during eclipse; and (3) using a tether for eccentricity pumping to correct perturbations in the orbit and as a means of adding energy to the orbit for boosting and orbital transfer.

Author N89-26878*# National Aeronautics and Space Administration. Langley Research Center, Hampton, VA. TWO-STAGE REUSABLE LAUNCH SYSTEM UTILIZING A WINGED CORE VEHICLE AND GLIDEBACK BOOSTERS lan O. MacConochie, James A. Martin, James S. Wood, and Miles O. Duquette (Joint Inst. for Advancement of Flight Sciences, Washington, DC.) Jul. 1989 21 p (NASA-TM-101513; NAS 1.15:101513) Avail: NTIS HC A03/MF A01 CSCL 22B

A near-term technology launch system is described in which

N89-26880*# Harris Corp., Melbourne, FL. ADVANCED MODULATION TECHNOLOGY DEVELOPMENT FOR EARTH STATION DEMODULATOR APPLICATIONS Final Report R. C. Davis, J. V. Wernlund, J. A. Gann, J. F. Roesch, T. Wright, and R. D. Crowley Jul. 1989 282 p (Contract NAS3-24681) (NASA-CR-185126; NAS 1.26:185126) Avail: NTIS HC A13/MF A01 CSCL 09F

The purpose of this contract was to develop a high rate (200 Mbps), bandwidth efficient, modulation format using low cost hardware, in 1990's technology. The modulation format chosen is 16-ary continuous phase frequency shift keying (CPFSK). The implementation of the modulation format uses a unique combination of a limiter/discriminator followed by an accumulator to determine transmitted phase. An important feature of the modulation scheme is the way coding is applied to efficiently gain back the performance lost by the close spacing of the phase points.

Author

18 SPACECRAFT DESIGN, TESTING AND

PERFORMANCE

9 p

Includes satellites; space platforms; space stations; spacecraft systems and components such as thermal and environmental controls; and attitude controls.

For life support systems see 54 Man/System Technology and Life Support. For related information see also 05 Aircraft Design, Testing and Performance, 39 Structural Mechanics, and 16 Space Transportation.

N89-26884# Instituto de Pesquisas Espaciais, Sao Jose dos
Campos (Brazil).
STUDY OF JITTER ON THE ATTITUDE MOTION OF A SPIN
STABILIZED SATELLITE WITH TWO FLEXIBLE SOLAR
PANELS (ESTUDO DO JITTER NA ATITUDE DE UM SATELITE
ARTIFICIAL COM PAINEIS FLEXIVEIS ESTABILIZADOS FOR
ROTACAO DUAL)
ljar MilagredaFonseca Dec. 1988

In PORTUGUESE; ENGLISH summary Presented at the 7th Brazilian Conference on automation, Sao Jose dos Campos, Brazil, 15-19 Aug. (INPE-4744-PRE/1418) Avail: NTIS HC A02/MF A01

The jitter on the attitude motion of a spin stabilized satellite with two flexible solar panels is studied. The panels vibration displacements are discretized by the assumed modes method to obtain a model described by ordinary differential equations only. The Lagrangian formulation was used to derive the equations of motion. These equations were integrated numerically and the results, presented in graphical form, show the effects of the rotor misalignment and of the flexibility of the solar panels on the attitude motion. The results also show that the Jitter specification for the satellite under study is satisfied.

Author

N89-26881# Office of Technology Assessment, Washington,
DC.
BIG DUMB BOOSTERS: A LOW-COST SPACE
TRANSPORTATION OPTION. AN OTA (OFFICE OF
TECHNOLOGY AND ASSESSMENT) BACKGROUND PAPER
Feb. 1989 35 p
(PB89-155196) Avail: NTIS HC A03/MF A01 CSCL 22B

A launch vehicle concept commonly known as the Big Dumb Booster, a concept that derives from efforts first made in the 1960s to minimize costs of space launch systems is described and examined. Some launch system analysts believe that the use of this concept, when applied to existing technology, could markedly reduce space transportation costs. Other analysts disagree. Low-cost space transportation is one of the keys to more effective exploration and exploitation of outer space. If space transportation costs were much lower, government agencies and firms with good ideas for using the space environment might be more willing to risk their investment capital. In this area of increased budget stringency, the high cost of space transportation has prompted analysts to examine a wide variety of ideas to reduce these costs.

GRA

N89-26885* # Teledyne Brown Engineering, Huntsville, AL. ADVANCED SOLAR OBSERVATORY (ASO) ACCOMMODATIONS REQUIREMENTS STUDY Final Report May 1989 330 p (Contract NAS8-37128) (NASA-CR-183718; NAS 1.26:183718; SP89-MSFC-3225) Avail: NTIS HC A15/MF A01 CSCL 22B

Results of an accommodations analysis for the Advanced Solar Observatory on Space Station Freedom are reported. Concepts for the High Resolution Telescope Cluster, Pinhole/Occulter Facility, and High Energy Cluster were developed which can be accommodated on Space Station Freedom. It is shown that workable accommodations concepts are possible. Areas of emphasis for the next stage of engineering development are identified.

Author

6 p

N89-26882 Department of the Air Force, Washington, DC.
LOW LOSS RADAR WINDOW FOR REENTRY VEHICLE Patent
Thomas M. Place, Jr., inventor (to AF) 22 Nov. 1988
Filed 10 Sep. 1987
(AD-D014045; US-Patent-4,786,548; US-Patent-Appl-SN-095062)
Avail: US Patent and Trademark Office CSCL 17/9

A radar window for a reentry vehicle where the window is made of a composite material of boron nitride, silica, and boron nitride yarn is described. Silica in the range of 2 to 18 percent is the preferred range in the composite material and the boron nitride matrix ranges from 5 to 15 percent.

GRA

N89-26886# Instituto de Pesquisas Espaciais, Sao Jose dos
Campos (Brazil). Dept. de Controle e Guiagem.
SATELLITE ATTITUDE CONTROL ANALYSIS AND DESIGN
USING LINEARIZING TRANSFORMATIONS
Antonio Felix Martin Neto Dec. 1988 7 p Presented at the
7th Congresso Brasileiro de Automatica, Sao Jose dos Campos,
Brazil Submitted for publication
(INPE-4772-PRE/1441) Avail: NTIS HC A02/MF A01

The problem of maneuvering a satellite relative to a desired noninertial reference is examined by using linearizing transformations and the concept of relative quaternion. The rigid body attitude equations are transformed into an equivalent linear form which considers the fact that the control purpose is to align the satellite with a moving reference frame (e.g., and Earth resource satellite). The equations so obtained are used to design control strategies for stabilizing the satellite attitude by linear techniques. This approach is more useful for algorithm design for on-board processors than the technique using inertial quaternion since the laws obtained are a function of the noninertial sensor outputs (sun sensors, earth sensors, etc.) and of the inertial sensor (gyroscopes) outputs.

Author

N89-26883# Optimization, Inc., Blacksburg, VA. CONTROL OF FLEXIBLE STRUCTURES Final Report, 1 Mar. 1986 - 30 Sep. 1988 John A. Burns, Eugene M. Cliff, H. J. Kelley, F. H. Lutze, and R. E. Miller Apr. 1989 151 p (Contract F04611-86-C-0008) (AD-A207520; K511-3; AL-TR-89-001) Avail: NTIS HC A08/MF A01 CSCL 13/13

This report summarizes the work done under contract FO 4611-86-C-0008. The principal goals were to develop state-space models and computational algorithms for control of beam and plate type structures, and, more generally, to increase the understanding of the basic problems associated with this development. The state-space approach is based on a distributed parameter model of the structure that includes the fundamental equations without modal truncation. The approach is to use basic physical principles to write down the governing partial differential equations, construct a state-space model from these governing equations, formulate the optimal control problem in terms of the state-space model, develop a convergent approximation scheme and conduct numerical experiments to test the method. GRA

N89-26887*# National Aeronautics and Space Administration.
Lewis Research Center, Cleveland, OH.
PHOTOVOLTAIC MODULE ON-ORBIT ASSEMBLY FOR SPACE
STATION FREEDOM
Thomas Sours, R. Lovely, and D. Clark 1989 10 p Presented
at the 24th Intersociety Energy Conversion Engineering Conference,
Washington, DC, 6-11 Aug. 1989; cosponsored by the IEEE, AIAA,
ANS, ASME, SAE, ACS, and AICHE
(NASA-TM-102297; E-4973; NAS 1.15:102297) Avail: NTIS HC
A02/MF A01 CSCL 22B

One of the elements of the Space Station Freedom power system is the Photovoltaic (PV) module. These modules will be assembled on-orbit during the assembly phase of the program. These modules will be assembled either from the shuttle orbiter or from the Mobile Servicing Center (MSC). The different types of assembly operations that will be used to assemble PV Modules are described.

Author

N89-26888*# National Aeronautics and Space Administration.
Langley Research Center, Hampton, VA.
ALTERNATE SPACE STATION FREEDOM CONFIGURATION
CONSIDERATIONS TO ACCOMMODATE SOLAR DYNAMIC
POWER
L. J. DeRyder, J. N. Cruz, M. L. Heck, B. P. Robertson, and P. A.
Troutman 28 May 1989 35 p
(NASA-TM-101603; NAS 1.15:101603) Avail: NTIS HC
A03/MF A01 CSCL 22B

The results of a technical audit of the Space Station Freedom Program conducted by the Program Director was announced in early 1989 and included a proposal to use solar dynamic power generation systems to provide primary electrical energy for orbital flight operations rather than photovoltaic solar array systems. To generate the current program baseline power of 75 kW, two or more solar concentrators approximately 50 feet in diameter would be required to replace four pairs of solar arrays whose rectangular blanket size is approximately 200 feet by 30 feet. The photovoltaic power system concept uses solar arrays to generate electricity that is stored in nickel-hydrogen batteries. The proposed concept uses the solar concentrator dishes to reflect and focus the Sun's energy to heat helium-xenon gas to drive electricity generating turbines. The purpose here is to consider the station configuration issues for incorporation of solar dynamic power system components. Key flight dynamic configuration geometry issues are addressed and an assembly sequence scenario is developed.

Author

Two modeling methods for the sloshing dynamics of liquids in partially filled tanks in a spinning spacecraft, the homogeneous vortex flow approximation and the boundary layer model, are presented. A modeling method used to implement liquid sloshing effects into the simulation program DISCOS for multibody systems is described. The computer programs for the analysis of sloshing effects and the extended simulation program DISCOSL were applied to the INTELSAT 6 spacecraft.

ESA N89-26892# Naval Postgraduate School, Monterey, CA. SPIN STABILIZATION OF THE ORION SATELLITE USING A THRUSTER ATTITUDE CONTROL SYSTEM WITH OPTIMAL CONTROL CONSIDERATIONS M.S. Thesis Janet L. Cunningham Mar. 1989 48 p (AD-A208484) Avail: NTIS HC A03/MF A01 CSCL 22/1

The controlled system is the ORION satellite spinning about its single axis of symmetry. Hydrazine thrusters are used as the control and are modeled by ideal, constant magnitude step functions. The system is normalized and driven from non-zero initial angular velocities of the two axes other than the spin axis to the final condition zero. The control profiles required to do this are determined based on a desired controller duty cycle. Adaption of the duty cycle changes the ratio of the time the thrusters are on (fuel use) and total time to completion of the evolution. A comparison between a single axis and a dual axis controller is presented. Simulation programs for the normalized system using a single axis controller simulation program, with each controller having a duty cycle of no more than 50 percent are developed. Operation of the system is optimized using a system cost function. An equation relating the controller duty cycle of the dual system to the fuel/time trade-off parameter of the system cost function is required. A nonlinear feedback control algorithm (function of attitude angle rates) is developed from iterations of the simulation, and a priori knowledge of the form of the control from optimal control theory. This numerical solution will allow system designers to incorporate a closed form state feedback control for minimum fuel/time strategies using the ORION satellite's onboard software.

GRA

N89-26889# Centre National d'Etudes Spatiales, Toulouse (France). Div. Techniques Vehicules. SPOT 1-2-3 PYROTECHNICS SYSTEM: DEFINITION REPO RT (PROJETS SPOTS 1, 2 ET 3. SYSTEME PYROTECHNIQUE. DOSSIER DE SYNTHESE DE DEFINITION) Jean-Pierre Bouloumie 2 Mar. 1989 67 p In FRENCH Original contains color illustrations (CNES-CT/DRT/TVE/IL-89-035; ETN-89-94808) AVAIL: NTIS HC A04/MF A01

Information on platform and payload pyrotechnical systems of Spot projects 1, 2, and 3 is presented. The technical specifications and definition sheets are summarized to describe the systems. The description includes separation of Ariane launcher, antenna S2 deployment, deblocking optical instruments, deployment of solar panels, and deblocking of the solar generator drive.

ESA

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N89-26890# Messerschmitt-Boelkow-Blohm G.m.b.H., Bremen (Germany, F.R.). COLUMBUS LOGISTICS PROGRAM Mike C. Attwood 1988 11 p Previously announced in IAA as A89-18314 (MBB-UO-0039/88-Pub; ETN-89-94617) Avail: NTIS HC A03/MF A01

The overall logistics scenario which reflects the design configurations and operational concepts as derived from the Columbus phase B activities and results of considerations of coherence with other European space elements, (e.g., Hermes). is described. The required resupply infrastructure, covering the range from provisions for off-line repair of units at a manufacturer's site via ground/orbit transport to on-orbit storage, until utilization in a servicing activity, are identified.

ESA

20 SPACECRAFT PROPULSION AND

POWER

Includes main propulsion systems and components, e.g., rocket engines; and spacecraft auxilliary power sources.

For related information see also 07 Aircraft Propulsion and Power, 28 Propellants and Fuels, 44 Energy Production and Conversion, and 15 Launch Vehicles and Space Vehicles.

N89-26891# Messerschmitt-Boelkow-Blohm G.m.b.H., Ottobrunn
(Germany, F.R.). Space Communications and Propulsion Systems
Div.
MODELING OF LIQUID SLOSHING EFFECTS IN MULTI BODY
SYSTEMS
K. Ebert 1988 7p
(MBB-UK-0024/88-Pub; ETN-89-94618) Avail: NTIS HC
A02/MF A01

N89-26893# National Aerospace Lab., Tokyo (Japan). LH SUB 2/LO SUB 2 ROCKET ENGINE WITH SOLAR ENERGY Takeshi Kanda, Yoshio Wakamatsu, and Akio Kanmuri Aug 1988 16 p In JAPANESE; ENGLISH summary (NAL-TR-992; ISSN-0389-4010) Avail: NTIS HC A03/MF A01

Solar energy is already used in the form of electric power generated by solar cells in satellites. And a solar thermal propulsion system, a kind of gas jet engine, was studied. Large expansion area ratio engines were studied to achieve high specific impulse, Isp. If external energy is added to an engine, Isp can be better than the theoretical Isp based on tank level enthalpy. Solar energy was used as the additional energy supplied to a rocket engine. The LH2/LO2 expander cycle engine with a solar heat exchanger was studied. This engine has two merits. One is the increase of Isp, which allows a greater payload to be carried by a rocket with the solar heat exchanger than by a normal rocket. The other merit is a cycle pressure level decrease. The Isp increase is realized for every engine and for every propellant, if additional energy is available.

Author

monitoring the operational health of the Space Shuttle Main Engine (SSME) are being evaluated. The ultimate goal is providing protection for the SSME as well as improving ground and flight test techniques. One scenario with some potential is measuring radio frequency (RF) emissions (if present) in the exhaust plume and correlating the data to engine health. An RF emissions detection system was therefore designed, the equipment leased, and the components integrated and checked out to conduct a quick-look investigation of RF emissions in the SSME exhaust plume. The system was installed on the A-1 Test Stand at Stennis Space Center, MS, and data were successfully acquired during SSME firings from May 3 to September 15, 1988. The experiments indicated that emitted radiation in the RF (20 to 470 MHz) spectrum definitely exists in the SSME exhaust plume, and is of such magnitude that it can be distinguished during the firing from background noise. Although additional efforts are necessary to assess the merit of this approach as a health monitoring technique, the potential is significant, and additional studies recommended.

Author

226 p.

are

N89-26894*# Lockheed Missiles and Space Co., Huntsville, AL.
Engineering Center.
SPACE SHUTTLE MAIN ENGINE STRUCTURAL ANALYSIS
AND DATA REDUCTION/EVALUATION. VOLUME 7: HIGH
PRESSURE FUEL TURBO-PUMP THIRD STAGE IMPELLER
ANALYSIS Final Report
Kirby V. Pool Apr. 1989

Original contains color illustrations (Contract NAS8-37282) (NASA-CR-183670; NAS 1.26:183670; LMSC-HEC-TR-F268584-Vol-7) Avail: NTIS HC A11/MF A01 CSCL 21H

This volume summarizes the analysis used to assess the structural life of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbo-Pump (HPFTP) Third Stage Impeller. This analysis was performed in three phases, all using the DIAL finite element code. The first phase was a static stress analysis to determine the mean (non-varying) stress and static margin of safety for the part. The loads involved were steady state pressure and centrifugal force due to spinning. The second phase of the analysis was a modal survey to determine the vibrational modes and natural frequencies of the impeller. The third phase was a dynamic response analysis to determine the alternating component of the stress due to time varying pressure impulses at the outlet (diffuser) side of the impeller. The results of the three phases of the analysis show that the Third Stage Impeller operates very near the upper limits of its capability at full power level (FPL) loading. The static loading alone creates stresses in some areas of the shroud which exceed the yield point of the material. Additional cyclic loading due to the dynamic force could lead to a significant reduction in the life of this part. The cyclic stresses determined in the dynamic response phase of this study are based on an assumption regarding the magnitude of the forcing function.

K.C.D.

N89-26897 Department of the Air Force, Washington, DC. CONTEMPORARY COMPOSITE POLAR BOSS Patent Hugh M. Reynolds, inventor (to AF) and Curt M. Kawabata, inventor (to AF) 28 Feb. 1989 9p Filed 16 Oct. 1987 (AD-D014074; US-Patent-4,807,531; US-Patent-Appl-SN-109557; US-Patent-Class-102-347) Avail: US Patent and Trademark Office CSCL 21/8

Composite polar bosses applied in rocket motor cases have very desirable characteristics. The lead time for a boss can be reduced from 6 to 12 months to 2 to 3 months. Weight savings for the boss are about 20 to 40 pct. A composite polar boss which attaches closures to solid fueled rocket motor cases is disclosed. This polar boss is a carbonized fabrication which sits within a circumferential indentation within the motor case, and has a threaded inner circumference which permits the closure to be attached thereto. The materials of the boss are selected to permit a service temperature of 350 F.

GRA

N89-26898# Aerospace Corp., El Segundo, CA. Aerophysics Lab. HYBRID ELECTRIC CHEMICAL PROPULSION James E. Pollard and Ronald B. Cohen 1 Apr. 1989 27 p (Contract F04701-85-C-0086) (AD-A207584; TR-0086(6930-03)-1; SD-TR-89-24) Avail: NTIS HC A03/MF A01 CSCL 21/8

A spacecraft propulsion concept that uses a combination of electrical energy and chemical energy for thrust is explored using thermodynamic modeling calculations. The essence of this concept consists of adding a carefully chosen amount of O2 or F2 oxidizer to the propellant flow of a conventional H2 electrothermal thruster. A general method is given for selecting the fuel - an oxidizer ratio so as to optimize the thruster performance for any set of mission constraints.

GRA

N89-26895*# National Aeronautics and Space Administration, Washington, DC. THE NASA LOW THRUST PROPULSION PROGRAM James R. Stone and Gary L. Bennett Jul. 1989 31 p Presented at the 25th Joint Propulsion Conference, Monterey, CA, 10-12 Jul. 1989; sponsored in part by AIAA, ASME, SAE, and ASEE (NASA-TM-102065; E-4822; NAS 1.15:102065; AIAA-89-2492) Avail: NTIS HC A03/MF A01 CSCL 10B

The NASA OAST opulsion, Power, and ergy Division supports a low thrust propulsion program aimed at providing high performance options for a broad range of near-term and far-term mission and vehicles. Low thrust propulsion has a major impact on the mission performance of essentially all spacecraft and vehicles. On-orbit lifetimes, payloads, and trip times are significantly impacted by low thrust propulsion performance and integration features for Earth-to-orbit (ETO) vehicles, Earth-orbit and planetary spacecraft, and large platforms in Earth orbit. Major emphases

on low thrust chemical propulsion, both storables and hydrogen/oxygen; low-power (auxiliary) electric arcjects and resistojets; and high-power (primary) electric propulsion, including ion, magnetoplasmadynamic (MPD), and electrodeless concepts. The major recent accomplishments of the program are presented and their impacts discussed.

Author

are

38 p

N89-26899*# Rockwell International Corp., Canoga Park, CA.
Rocketdyne Div.
OTVE COMBUSTOR WALL CONDITION MONITORING Final
Report, Nov. 1986 - Sep. 1988
Brian Szemenyei, Robert S. Nelson, and S. Barkhoudarian Aug.
1989

Presented at the 24th AIAA/ASME/SAE/ASEE Joint Propulsion Conference on Improved Maintainability of Space-Based Reuseable Rocket Engines, Boston, MA, Jul. 1988 (Contract NAS3-23773) (NASA-CR-182275; NAS 1.26:182275; RI/RD89-212) Avail: NTIS HC A03/MF A01 CSCL 21H

Conventional ultrasonics, eddy current, and electromagnetic acoustic transduction (EMAT) technologies were evaluated to determine their capability of measuring wall thickness/wear of individual cooling channels in test specimens simulating conditions in the throat region of an OTVE combustion chamber liner. Quantitative results are presented for the eddy current technology, which was shown to measure up to the optimum 20-mil wall thickness with near single channel resolution. Additional results

N89-26896*# Sverdrup Technology, Inc., Bay Saint Louis, MS.
SPACE SHUTTLE MAIN ENGINE RADIO FREQUENCY
EMISSIONS Final Report, May - Sep. 1988
A. W. Rester, E. L. Valenti, and L. R. Smith Nov. 1988 23 p
(NASA-CR-184707; NAS 1.26:184707) Avail: NTIS HC
A03/MF A01 CSCL 21H

Several approaches to develop a diagnostics system for

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