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NASA NTRS1983

Computing the projected area of a spacecraft solar array with matrices

Matrix methods for computing the projected area of a solar array as a function of the rotational position of a spacecraft and of array position on a spacecraft are presented. Formulas are derived which provide the optimum solar array pitch, cant and tracking angles for a given spacecraft configuration and orbit. These formulas are general and applicable to many spacecraft. Formulas are also provided for determining the energy output from an array for a given orbit and the resultant energy available to spacecraft loads. Results are simply obtained and realizable with a hand calculator. The methods above can be extended to the case of computer analysis of solar array shadowing. How this can be done is outlined and results are presented from a spacecraft study program.

E. M. Gaddy↗

NASA NTRS2001

Thermal Design Overview of the Mars Exploration Rover Project

This slide presentation reviews the thermal design for the Mars exploration rover project. It includes information on the spacecraft configuration, the cruise scenario, landing scenario, instrument package, thermal environment, and spacecraft schematics.

Mars Exploration Rovers↗

NASA NTRS1970

HELIOS Configuration for Close Sun Flyby

This report is the documented result of an Internal study effort to examine possible spacecraft configurations for flight to within 0.20 AU of the sun. The objective of this study was to optimize the HELIOS spacecraft design for flights closer than 0.3 AU without jeopardizing the HELIOS Project costs or schedule. The results of the study meet this objective. In addition to this written report, an oral presentation has been prepared for the Second HELIOS Joint Working Group Meeting at GSFC on 27 April 1970. In the performance of this study, Messerschmitt-Bolkow-Blohm GMBH was assisted by the General Electric Company, Missile and Space Division, Valley Forge, Pennsylvania.

Lex, J.↗

NASA NTRS1987

A high-capacity aeronautical mobile satellite system

This paper describes a conceptual system design for a satellite-based aeronautical safety communications system capable of serving both general aviation aircraft and commercial aviation aircraft in the contiguous U.S. in the mid-1990s. The space segment is described, including satellite locations and coverage, spacecraft configuration, eclipse capability and stationkeeping, transponder design, and mass and power. The spacecraft mass and power budgets are given. The air mobile terminals, ground segment, and frequency plan and channelization are discussed, and the data rate, modulation/demodulation/coding, and channel spacing are considered. The message format, frequency control, system capacity, and system sensitivity are discussed.

Sue, M. K.↗

NASA NTRS1972

Applications technology satellites advanced mission study

Three spacecraft configurations were designed for operation as a high powered synchronous communications satellite. Each spacecraft includes a 1 kw TWT and a 2 kw Klystron power amplifier feeding an antenna with multiple shaped beams. One of the spacecraft is designed to be boosted by a Thor-Delta launch vehicle and raised to synchronous orbit with electric propulsion. The other two are inserted into a elliptical transfer orbit with an Atlas Centaur and injected into final orbit with an apogee kick motor. Advanced technologies employed in the several configurations include tubes with multiple stage collectors radiating directly to space, multiple-contoured beam antennas, high voltage rollout solar cell arrays with integral power conditioning, electric propulsion for orbit raising and on-station attitude control and station-keeping, and liquid metal slip rings.

Gould, L. M.↗

NASA NTRS1972

Applications technology satellite advanced missions study, volume 1

Four different spacecraft configurations were developed for geostationary service as a high power communications satellite. The first configuration is a Thor-Delta launch into a low orbit with a spiral ascent to synchronous altitude by ion engine propulsion. The spacecraft is earth oriented with rotating solar arrays. Configuration 2 is a direct injection Atlas/Centaur/Burner II vehicle which when in orbit is sun-oriented with a rotating transponder tower. Configurations 3 and 4 are Titan IIIC launches, and are therefore larger and heavier than Configuration 2. They are both sun-oriented, with rotating transponder towers and are directly injected into orbit. Technology discussed includes high power (up to 2 kW) transmitters with collectors radiating heat directly into space, and contoured antenna patterns designed to illuminate particular earth regions. There is also a review of potential users of the services which can be performed by this type satellite in such areas as information networking, public broadcasting and educational television.

Robinson, D. L.↗

NASA NTRS1973

Microbiological sampling of spacecraft cabling, antennas, solar panels and thermal blankets

Sampling procedures and techniques described resulted from various flight project microbiological monitoring programs of unmanned planetary spacecraft. Concurrent with development of these procedures, compatibility evaluations were effected with the cognizant spacecraft subsystem engineers to assure that degradation factors would not be induced during the monitoring program. Of significance were those areas of the spacecraft configuration for which special handling precautions and/or nonstandard sample gathering techniques were evolved. These spacecraft component areas were: cabling, high gain antenna, solar panels, and thermal blankets. The compilation of these techniques provides a historical reference for both the qualification and quantification of sampling parameters as applied to the Mariner Spacecraft of the late 1960's and early 1970's.

Koukol, R. C.↗

NASA NTRS1991

Hoop column soil moisture spacecraft in low Earth orbit for global change monitoring

A subset of the total Global Change Technology Initiative instruments are required to be in low Earth, sunsynchronous orbits. There is one instrument, however, that requires its own specialized spacecraft; the Soil Moisture Microwave Radiometer (SMMR). The characteristic structure of the instrument is the 118 m hoop column support structure. The hoop is supported by an axially placed column. Tension cables support and shape an electromagnetically reflective mesh surface. The instrument is capable of detecting frequencies in the 1.4 GHz range (Soil Moisture and Sea Salinity). Three apertures are used to reduce the degree of paraboloid offset and improve the beam quality. The spacecraft configuration is determined by the instrument support requirements and the requirement that it can fit into the Titan IV cargo bay. The configuration is derived by cross referencing the instrument performance requirements with the performance of the spacecraft. The spacecraft design is similar with the Multi-mission Modular Spacecraft in terms of size and packaging. A description of the spacecraft's features will yield a summary of the technologies needed for the SMMR spacecraft.

Ferebee, Melvin J., Jr.↗

NASA NTRS2008

Potential of a New Lunar Surface Radiator Concept for Hot Lunar Thermal Environments

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft s vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed approx.325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided. A recent study of vertically oriented radiator configurations subjected to lunar noon thermal environments led to the discovery of a novel radiator concept that yielded positive heat rejection at lower fluid temperatures. This radiator configuration, called the Upright Lunar Terrain Radiator Assembly (ULTRA), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of the ULTRA is the absence of louvers or other moving parts and its simple geometry. Analysis of the ULTRA for a lunar extravehicular activity (EVA) portable life support system (PLSS) is shown to provide moderate heat rejection, on average, at all solar incident angles assuming an average radiator temperature of 294 K, whereas prior concepts exhibited insignificant heat rejection or heat absorption at higher incident angles. The performance of the ULTRA for a lunar lander is also discussed and compared to the performance of a vertically oriented, flat panel radiator at various lunar latitudes.

Ochoa, Dustin A.↗

NASA NTRS2013

Optimized Radiator Geometries for Hot Lunar Thermal Environments

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft's vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed approximately 325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided. A recent study of vertically oriented radiator configurations subjected to lunar noon thermal environments led to the discovery of a novel radiator concept that yielded positive heat rejection at lower fluid temperatures. This radiator configuration, called the Intense Thermal Infrared Reflector (ITIR), has exhibited superior performance to all previously analyzed concepts in terms of heat rejection in the lunar noon thermal environment. A key benefit of ITIR is the absence of louvers or other moving parts and its simple geometry (no parabolic shapes). ITIR consists of a specularly reflective shielding surface and a diffuse radiating surface joined to form a horizontally oriented V-shape (shielding surface on top). The point of intersection of these surfaces is defined by two angles, those which define the tilt of each surface with respect to the local horizontal. The optimum set of these angles is determined on a case-by-case basis. The idea assumes minimal conductive heat transfer between shielding and radiating surfaces, and a practical design would likely stack sets of these surfaces on top of one another to reduce radiator thickness.

Ochoa, Dustin↗

NASA NTRS1985

Spacecraft (Mobile Satellite) configuration design study

The relative costs to procure and operate a two-satellite mobile satellite system designed to operate either in the UHF band of the L Band, and with several antenna diameter options in each frequency band was investigated. As configured, the size of the spacecraft is limited to the current RCA Series 4000 Geosynchronous Communications Spacecraft bus, which spans the range from 4000 to 5800 pounds in the transfer orbit. The Series 4000 bus forms the basis around which the Mobile Satellite transponder and associated antennas were appended. Although the resultant configuration has little outward resemblance to the present Series 4000 microwave communications spacecraft, the structure, attitude control, thermal, power, and command and control subsystems of the Series 4000 spacecraft are all adapted to support the Mobile Satellite mission.

Source record↗

NASA NTRS1993

Structural Mode Identification of Galileo Spacecraft from Flight Data

The Galileo spacecraft was launched in 1989 and is on its way to explore Jupiter. The dual-spin configured spacecraft will orbit Jupiter to conduct scientific investigation of the planet and its satellites. A probe will be released prior to Jupiter orbit insertion and will follow an impact trajectory for atmospheric investigation. A sketch of the Galileo spacecraft is shown in Fig. 1. The scan platform of the Galileo spacecraft is attached to a flexible stator structure that is in turn attached to the rotor (spinning portion) of the spacecraft.

Glenn A. Macala↗

NASA NTRS1980

Technologies involved in configuring an advanced earth-to-orbit transport for low structural mass

A tradeoff study for a single-stage, earth-to-orbit transport spacecraft configured with low structural mass as the prime objective is presented. Among the major design elements affecting vehicle mass are: maneuverability and re-entry heating, the compatibility of an irregular cargo bay geometry to future mission requirements, the fabrication and assembly of large honeycomb structure fuselage sections, and the development of LOX/LH2 engine extendable nozzles and dual-propulsion schemes. It is shown that vehicle geometry and materials and structural technologies will be the critical areas in the development of such a system.

Macconochie, I. O.↗

NASA NTRS1975

Spacecraft attitude control for a solar electric geosynchronous transfer mission

A study of the Attitude Control System (ACS) is made for a solar electric propulsion geosynchronous transfer mission. The basic mission considered is spacecraft injection into a low altitude, inclined orbit followed by low thrust orbit changing to achieve geosynchronous orbit. Because of the extended thrusting time, the mission performance is a strong function of the attitude control system. Two attitude control system design options for an example mission evolve from consideration of the spacecraft configuration, the environmental disturbances, and the probable ACS modes of operation. The impact of these design options on other spacecraft subsystems is discussed. The factors which must be considered in determining the ACS actuation and sensing subsystems are discussed. The effects of the actuation and sensing subsystems on the mission performance are also considered.

Leroy, B. E.↗

NASA NTRS1975

Spacecraft attitude control for a solar electric geosynchronous transfer mission

A study of the Attitude Control System (ACS) is made for a solar electric propulsion geosynchronous transfer mission. The basic mission considered is spacecraft injection into a low altitude, inclined orbit followed by low thrust orbit changing to achieve geosynchronous orbit. Because of the extended thrusting time, the mission performance is a strong function of the attitude control system. Two attitude control system design options for an example mission evolve from consideration of the spacecraft configuration, the environmental disturbances, and the probable ACS modes of operation. The impact of these design options on other spacecraft subsystems is discussed. The paper presents a discussion of the factors which must be considered in determining the ACS actuation and sensing subsystems. The effects of the actuation and sensing subsystems on the mission performance are also considered.

Leroy, B. E.↗

NASA NTRS1975

Summary report on the Viking 1975 DSN telecommunications compatibility test program

The system design tests and test results that provided the basis for establishment of telecommunications design between the DSN and Viking 1975 were described. The Viking 1975/DSN Telecommunications Compatibility Test Program Consisted of three phases: subsystem design, system design, and system verification tests which were performed at JPL and at the Air Force Eastern Test Range and Kennedy Space Center complexes. Subsystem design tests were performed with the Viking Orbiter (VO) and the Viking Lander (VL) during 1973. System design compatibility tests were performed with the Viking Proof Test Orbiter, Viking Spacecraft Test Lander, and a multiple Viking spacecraft configuration during the summer of 1974. System verification tests were performed with the Viking Orbiter, Viking Lander and Viking spacecraft during the spring and summer of 1975.

A. I. Bryan↗

NASA NTRS1988

ISE structural dynamic experiments

The topics are presented in viewgraph form and include the following: directed energy systems - vibration issue; Neutral Particle Beam Integrated Space Experiment (NPB-ISE) opportunity/study objective; vibration sources/study plan; NPB-ISE spacecraft configuration; baseline slew analysis and results; modal contributions; fundamental pitch mode; vibration reduction approaches; peak residual vibration; NPB-ISE spacecraft slew experiment; goodbye ISE - hello Zenith Star Program.

Lock, Malcolm H.↗

NASA NTRSDate not supplied

Automated End-to-End Spacecraft Connectivity Across Diverse Links

An increasing variety of communications services are available to space missions. Yet varying standards between providers hinder adoption due to the complexity of managing many configurations. We present a software framework to automatically establish end-to-end communications during contacts with a provider by configuring spacecraft protocols at the physical, link, and network layers. Underlying protocols and routes are abstracted, allowing the user to simply send data to a destination with the framework ensuring its delivery. We evaluate a full implementation of the framework in laboratory experiments conducted on an emulated communications testbed. These tests demonstrate data delivery across three different services with rapid (<20s) reconfiguration as the spacecraft transitions between providers.

satellite communication↗