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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 NTRS1984

Effects of the motions of a platform-mounted payload on the nutational stability of a dual-spin spacecraft - A Galileo case study

In the presence of spacecraft nutation, tracking of an inertially fixed target by the Galileo scan platform requires the use of two control loops to move the scan platform and the stator in such a way as to compensate for spacecraft motion. The effect of these control loops on spacecraft nutational stability is examined using an eigenvalue analysis approach as well as several computer analysis packages. It was found that the actions of these control loops tend to drive nutation to the point of neutralizing, and even overpowering the damping actions of the spacecraft nutation damper, for pointing directions close to the spacecraft's poles. Stable and unstable zones are mapped out for two sets of spacecraft mass properties, and contributions of rotor asymmetry and stator flexibility are also discussed.

Man, G. K.↗

NASA NTRS1984

A study of structural concepts for ultralightweight spacecraft

Structural concepts for ultralightweight spacecraft were studied. Concepts for ultralightweight space structures were identified and the validity of heir potential application in advanced spacecraft was assessed. The following topics were investigated: (1) membrane wrinkling under pretensioning; (2) load-carrying capability of pressurized tubes; (3) equilibrium of a precompressed rim; (4) design of an inflated reflector spacecraft; (5) general instability of a rim; and (6) structural analysis of a pressurized isotensoid column. The design approaches for a paraboloidal reflector spacecraft included a spin-stiffened design, both inflated and truss central columns, and to include both deep truss and rim-stiffened geodesic designs. The spinning spacecraft analysis is included, and the two truss designs are covered. The performances of four different approaches to the structural design of a paraboloidal reflector spacecraft are compared. The spinning and inflated configurations result in very low total masses and some concerns about their performance due to unresolved questions about dynamic stability and lifetimes, respectively.

Miller, R. K.↗

NASA NTRS1985

Designing a micro-spacecraft

Planetary spacecraft design could move toward less complex probes which would cost less then previous highly instrumented missions. The goal then becomes to fly more frequent missions and use commercial, proven hardware to ameliorate development costs. A commonality would be kept in place from spacecraft to spacecraft, with upgrades being introduced only to meet specific objectives or take advantage of advances in commercial hardware. Mission costs are in large part determined by spacecraft mass, so instrumentation must be miniaturized, i.e., the concept of a micro-satellite. A design study for the Cosimi project, which would feature placing a spacecraft on the far side of the solar corona to broadcast radio signals to earth, demonstrates the feasibility of a 20 cm diam rocket and integrated instruments for performing low-cost solar physics experiments. It is concluded, however, that current program start-ups will continue to maximize the mass and instrumentation of spacecraft.

Burke, J. D.↗

NASA NTRS1985

An historical overview of NASA manned spacecraft and their crew stations

The first generation of manned spacecraft demonstrated all the types of mission activity planned for the future. Development of these spacecraft and execution of their missions have provided the experience base for the design of current and future spacecraft systems. This paper summarizes the crew station characteristics of the Mercury, Gemini, Apollo, Skylab, Apollo-Soyuz and Space Shuttle spacecraft and the design strategies that dictated those characteristics. Also included are some comparisons with current and planned spacecraft and an assessment of the lessons learned as they might be applied in future spacecraft designs.

Loftus, J. P.↗

NASA NTRS1985

An assessment of the impact of spacecraft glow on the Hubble space telescope. Summary of existing observations and theory

Visible spacecraft glow was first observed on the Atmospheric Explorer spacecraft (AE-E) and studied in some detail with the Visible Airglow Experiment (VAE). The AE-E was a spin-stabilized spacecraft without thrusters at an altitude of 140 to 280 km. The VAE contained six visible wavelength photometers that measured a glow spectrum which: (1) rose steeply in the red, (2) decreased with a cos cubed PH1 dependence from pointing into the ram direction of the spacecraft orbital motion, and (3) decreased in intensity with increasing altitude with the same dependence as the measured atomic oxygen number atmospheric density (O) and not with the measured molecular nitrogen density (N sub 2). It is proposed that the glow is produced by chemical reactions on the spacecraft surface as it sweeps through the atmospheric O, with roughly 5-8 eV per O atom available for excitation from the orbital motion of the spacecraft. This glow may in principal be produced by any of a number of species, including molecular band emission from OH, NO, and NO2. An attempt is made to scale the observed glow to the Hubble space telescope.

Clarke, J. T.↗

NASA NTRS1985

Space power systems - 'Spacecraft 2000'

The program 'Spacecraft 2000' has the objective to identify critical, high-payoff, potential spacecraft technologies, taking into account the formulation, advocation, and the management of the requisite technology development programs. This program represents a joint NASA-industry program. The technology areas addressed by 'Spacecraft 2000' are related to spacecraft power/energy storage, thermal control/thermal management, power management and distribution, autonomous operation-control, on-board system integration, spacecraft environmental interactions, secondary propulsion, communications technologies, a total system response approach, and system-subsystem technology verification. The expected benefits of a development of advanced technologies include decreased spacecraft bus system weights, decreased mission costs, increased reliability/lifetimes, and increased operational flexibility.

Faymon, K. A.↗

NASA NTRS1986

Spacecraft 2000

The Spacecraft 2000 program has been conceived by NASA with the objective to maintain the U.S. in a strong technical position in the space arena. The program is to involve a cooperative NASA/industry effort with long-range technology development aims. The technology areas addressed are related to spacecraft power, spacecraft energy storage, power management and distribution, autonomous operation-control, onboard systems integration, secondary propulsion, communications technologies, and systems/subsystems technology verification. Aspects of Spacecraft 2000 interactions with industry are discussed, taking into account visitations of NASA personnel to selected industrial concerns, most critical concerns (general), major technical concern, technology drivers for future spacecraft power systems, technology considerations for secondary propulsion, and questions of commercialization. Plans related to a Spacecraft 2000 program formulation are also considered.

Faymon, K. A.↗

NASA NTRS1986

Orbital spacecraft resupply technology

The resupplying of orbital spacecraft using the Space Shuttle, Orbital Maneuvering Vehicle, Orbital Transfer Vehicle or a depot supply at a Space Station is studied. The governing factor in fluid resupply designs is the system size with respect to fluid resupply quantities. Spacecraft propellant management for tankage via diaphragm or surface tension configurations is examined. The capabilities, operation, and application of adiabatic ullage compression, ullage exchange, vent/fill/repressurize, and drain/vent/no-vent fill/repressurize, which are proposed transfer methods for spacecraft utilizing tankage configurations, are described. Selection of the appropriate resupply method is dependent on the spacecraft design features. Hydrazine adiabatic compression/detonation, liquid-free vapor venting to prevent freezing, and a method for no-vent liquid filling are analyzed. Various procedures for accurate measurements of propellant mass in low gravity are evaluated; a system of flowmeters with a PVT system was selected as the pressurant solubility and quantity gaging technique. Monopropellant and bipropellant orbital spacecraft consumable resupply system tanks which resupply 3000 lb of hydrazine and 7000 lb of MMH/NTO to spacecraft on orbit are presented.

Eberhardt, R. N.↗

NASA NTRS1987

Ka-band (32 GHz) spacecraft development plan

A road map for the development of a protoflight 32 GHz spacecraft solid state transmitter is given. The major milestones include the development of device and component technology required for use in the spaceborne experimental and operational transmitter systems. Two experimental spacecraft transmitter systems are envisioned: first, a low power beacon, to determine the performance of a 32 GHz downlink communications system; and second, an array feed, to further verify the results of the first experiment and service as a test bed for technology required for an operational system. The first experiment was proposed to NASA Headquarters for flight aboard the Mars Observer spacecraft with spacecraft integration in early 1989. The second is to be available for integration aboard a spacecraft such as the Comet Rendezvous Asteroid Flyby (CRAF) mission in the 1990 time frame. These experimental systems are to lead to the development of a protoflight transmitter for subsequent spacecraft integration in 1992, the time frame of the Cassini mission to Saturn.

Riley, A. L.↗

NASA NTRS1986

A unified approach to computer analysis and modeling of spacecraft environmental interactions

A new, coordinated, unified approach to the development of spacecraft plasma interaction models is proposed. The objective is to eliminate the unnecessary duplicative work in order to allow researchers to concentrate on the scientific aspects. By streamlining the developmental process, the interchange between theories and experimentalists is enhanced, and the transfer of technology to the spacecraft engineering community is faster. This approach is called the UNIfied Spacecraft Interaction Model (UNISIM). UNISIM is a coordinated system of software, hardware, and specifications. It is a tool for modeling and analyzing spacecraft interactions. It will be used to design experiments, to interpret results of experiments, and to aid in future spacecraft design. It breaks a Spacecraft Ineraction analysis into several modules. Each module will perform an analysis for some physical process, using phenomenology and algorithms which are well documented and have been subject to review. This system and its characteristics are discussed.

Katz, I.↗

NASA NTRS1987

The coupled dynamics of fluids and spacecraft in low gravity and low gravity fluid measurement

The very large mass fraction of liquids stored on broad current and future generation spacecraft has made critical the technologies of describing the fluid-spacecraft dynamics and measuring or gauging the fluid. Combined efforts in these areas are described, and preliminary results are presented. The coupled dynamics of fluids and spacecraft in low gravity study is characterizing the parametric behavior of fluid-spacecraft systems in which interaction between the fluid and spacecraft dynamics is encountered. Particular emphasis is given to the importance of nonlinear fluid free surface phenomena to the coupled dynamics. An experimental apparatus has been developed for demonstrating a coupled fluid-spacecraft system. In these experiments, slosh force signals are fed back to a model tank actuator through a tunable analog second order integration circuit. In this manner, the tank motion is coupled to the resulting slosh force. Results are being obtained in 1-g and in low-g (on the NASA KC-135) using dynamic systems nondimensionally identical except for the Bond numbers.

Hansman, R. John↗

NASA NTRS1987

Collision detection for spacecraft proximity operations

Collision Detection for Spacecraft Proximity Operations This thesis describes the development of a new collision detection algorithm to be used when two spacecraft are operating in the same vicinity. The two spacecraft are modelled as unions of convex polyhedra, where the polyhedron resulting from the union may be either convex or nonconvex. The relative motion of the two spacecraft is assumed to be such that one vehicle is moving with constant linear and angular velocity with respect to the other. The algorithm determines if a collision is possible and, if so, predicts the time when the collision will take place. The theoretical basis for the new collision detection algorithm is the C-function formulation of the configuration space approach recently introduced by researchers in robotics. Three different types of C-functions are defined that model the contacts between the vertices, edges, and faces of the polyhedra representing the two spacecraft. These C-functions are used to formulate three "collision" conditions. The first of these conditions limits the points representing potential collisions to the zeros of the C-functions. The new algorithm is fundamentally a search for the smallest zero of any C-function that satisfies the second and third collision conditions. The C-functions are shown to be transcendental functions of time for the assumed trajectory of the moving spacecraft. The zeros of these functions cannot be expressed in dosed form. Therefore, numerical search procedures are developed to find aLl of the zeros of a C-function in specified bounded intervals of time. These bounded intervals of time are found by examining the second and third collision conditions. The capabilities of the new algorithm are demonstrated for several example cases. These include examples of collisions determined by zeros of each of the three different types of C-functions. In addition to predicting the time of first contact of the polyhedra, the algorithm identifies the features of the two polyhedra that are touching at this time. The new collision detection algorithm is the first such algorithm that is capable of solving the collision detection problem exactly for the case where the moving object has constant linear and angular velocities. This is a significant improvement on previous collision detection algorithms described in the literature. In particular, the ability to handle constant angular velocity represents a more realistic type of rotational motion than those which have been used in other algorithms.

Robin M Vaughan↗

NASA NTRS1987

Development of a standard connector for orbital replacement units for serviceable spacecraft

The current trend for spacecraft to be serviceable and repairable in orbit has led to a modular approach to satellite subsystem design. Spacecraft equipment, such as sensors, tape recorders, computers, transponders, batteries, etc., housed in discrete modular units, (Orbital Replacement Units(ORUs)) can be attached and detached to the spacecraft as needed. The interface between the ORU and the spacecraft is crucial. The projected use of robotics and the need for a common mechanism capable of performing several functions puts many constraints on the design of the interface. Such an interface mechanisms was designed and developed called the Standard Interface Connector (SIC) that mates the ORU to the spacecraft. The SIC also provides for the flow of fluids, data, and power between the module and spacecraft. The baseline design presented can be configured to provide various attachment schemes. Tests on SIC models have demonstrated the functionality of the design and its compatibility with current robotics.

Heath, Ellen F.↗

NASA NTRS1988

Conceptual Spacecraft-Guidance Algorithm

Required weight of spacecraft minimized. Report describes conceptual algorithm for guidance of spacecraft launched from surface of Mars. Spacecraft to carry canister of specimens from surface to another spacecraft in orbit about Mars; second spacecraft then to carry canister back to Earth. Algorithm sufficiently general to be adaptable to prediction/correction algorithms for other spacecraft configurations.

Mccormick, Bernell R.↗

NASA NTRS1989

SHARP: A multi-mission AI system for spacecraft telemetry monitoring and diagnosis

The Spacecraft Health Automated Reasoning Prototype (SHARP) is a system designed to demonstrate automated health and status analysis for multi-mission spacecraft and ground data systems operations. Telecommunications link analysis of the Voyager II spacecraft is the initial focus for the SHARP system demonstration which will occur during Voyager's encounter with the planet Neptune in August, 1989, in parallel with real-time Voyager operations. The SHARP system combines conventional computer science methodologies with artificial intelligence techniques to produce an effective method for detecting and analyzing potential spacecraft and ground systems problems. The system performs real-time analysis of spacecraft and other related telemetry, and is also capable of examining data in historical context. A brief introduction is given to the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory. The current method of operation for monitoring the Voyager Telecommunications subsystem is described, and the difficulties associated with the existing technology are highlighted. The approach taken in the SHARP system to overcome the current limitations is also described, as well as both the conventional and artificial intelligence solutions developed in SHARP.

Lawson, Denise L.↗

NASA NTRS1989

Active rendezvous between a low-earth orbit user spacecraft and the Space Transportation System (STS) shuttle

Active rendezvous of an unmanned spacecraft with the Space Transportation System (STS) Shuttle is considered. The various operational constraints facing both the maneuvering spacecraft and the Shuttle during such a rendezvous sequence are discussed. Specifically, the actively rendezvousing user spacecraft must arrive in the generic Shuttle control box at a specified time after Shuttle launch. In so doing it must at no point violate Shuttle separation requirements. In addition, the spacecraft must be able to initiate the transfer sequence from any point in its orbit. The four-burn rendezvous sequence incorporating two Hohmann transfers and an intermediate phasing orbit as a low-energy solution satisfying the above requirements are discussed. The general characteristics of the four-burn sequence are discussed, with emphasis placed on phase orbit altitude and delta-velocity requirements. The planning and execution of such a sequence in the operational environment are then considered. Factor crucial in maintaining the safety of both spacecraft, such as spacecraft separation and contingency analysis, are considered in detail.

Hooper, H. L.↗

NASA NTRS1989

SHARP: A multi-mission artificial intelligence system for spacecraft telemetry monitoring and diagnosis

The Spacecraft Health Automated Reasoning Prototype (SHARP) is a system designed to demonstrate automated health and status analysis for multi-mission spacecraft and ground data systems operations. Telecommunications link analysis of the Voyager 2 spacecraft is the initial focus for the SHARP system demonstration which will occur during Voyager's encounter with the planet Neptune in August, 1989, in parallel with real time Voyager operations. The SHARP system combines conventional computer science methodologies with artificial intelligence techniques to produce an effective method for detecting and analyzing potential spacecraft and ground systems problems. The system performs real time analysis of spacecraft and other related telemetry, and is also capable of examining data in historical context. A brief introduction is given to the spacecraft and ground systems monitoring process at the Jet Propulsion Laboratory. The current method of operation for monitoring the Voyager Telecommunications subsystem is described, and the difficulties associated with the existing technology are highlighted. The approach taken in the SHARP system to overcome the current limitations is also described, as well as both the conventional and artificial intelligence solutions developed in SHARP.

Lawson, Denise L.↗