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The 'Baykonur' cosmodrome, its functions, operations, and services are described in considerable detail. The launch complex, launching pads, launch structures, launchers with cable masts and propellant loading towers, are included. The sequence of all phases of rocket assembly and preparations for launch are depicted. Prelaunch procedures and the launch itself are described.
Test data on the effect of the interaction between the space shuttle orbiter rear-mounted reaction control system jets and the flow over the vehicle in the atmosphere are analyzed with particular attention to yaw angle effects. Masted tank-plus-orbiter jet interactions resulting from combinations of forward and rear-mounted reaction control jets being fired prior to external tank staging are also documented. The primary correlating parameter used for the mated data is a single-jet momentum ratioed to free-stream momentum. An analytic model for computing reaction control system interactions for all control combinations used for mated vehicle coast is included.
The feasibility of modifying the solar electric propulsion (SEP) 66 watt/kilogram, 12.5 kilowatt solar array blanket design to incorporate ultra-low mass blanket technology and to generate conceptual design data by modifying the SEP solar array design to 17.5kW power output was performed. Five modified designs were developed, which substituted present SEP solar array design components with one or more of 50 micron thick solar cells, 75 micron cell coverglasses, and a different blanket substrate developed by GE. A parametric analysis was performed to determine the solar array mast least weight and blanket tension required to maintain a minimum natural frequency of 0.04 Hz. The solar array wing assembly weights and power outputs were calculated, and preliminary cost estimates for flight hardware development, fabrication and qualification were made for each case studied.
The distinctive feature of a tilt-rotor aircraft is that the pilot can change the rotor mast angles to go from a helicopter configuration for take-off and landing to an airplane configuration for high cruise speeds and vice-versa. An approach path for such an aircraft is proposed and the logic required to fly along this path in the presence of wind is determined. The main contribution of this work is an efficient and, to my knowledge, new method for generating the nominal state and control histories taking into account an estimate of the mean wind velocity and direction. The method requires the solution of algebraic (mostly linear) equations to generate a 'universal nominal', and feedforward and feedback gains. Then, in flight the additional state and control corrections due to deviation in descent rate, deceleration, and flight in a steady wind are obtained by multiplying simple precalculated functions of time by descent rate, deceleration or sine and cosine components of the mean wind vector. Simulations of approach flights for different wind conditions, assuming perfect state information in the feedback signal, indicated satisfactory performance.
All-mechanical antenna deployment system operates by single cable tensioned by electrically driven drum. Device is comprised of set of pulleys fixed to telescoping antenna mast, ratchet which prevents premature antenna retraction, and special latch which holds antenna in retracted position.
The Voyager spacecraft magnetometer experiment utilizes two sensors on a deployable boom. The boom is an Astromast. The implementation of the Astromast into the Voyager design is described along with the hardware used to hold, latch, and deploy the mast and the tests to demonstrate damping, deployment, and alignments. Several problems encountered are discussed and their solutions are given. Flight deployment and preliminary alignment results are presented. Finally, the design is evaluated in retrospect.
This paper describes the test program of a lightweight 25 kW solar array for solar electric propulsion. A full-scale development wing was made of aluminum with the containment box cover of graphite-epoxy, while the flight design array wing uses a graphite-epoxy structure. The full-scale continuous longeron array extension mast was tested for performance on a water table, and the full scale wing was functionally tested to demonstrate automatic containment box unlocking, wing extension, and retraction, blanket tensioning, and automatic application of blanket preload. The wing was then tested to the Shuttle acoustic environment, followed by a thermal/vacuum test in which the wing was extended and retracted at high and low temperature. Finally, the wing was tested in vibration with sine and random vibration environments.
The laser-one hazard detector system, used on the Rensselaer Mars rover, is reviewed briefly with respect to the hardware subsystems, the operation, and the results obtained. A multidetector scanning system was designed to improve on the original system. Interactive support software was designed and programmed to implement real time control of the rover or platform with the elevation scanning mast. The formats of both the raw data and the post-run data files were selected. In addition, the interface requirements were selected and some initial hardware-software testing was completed.
The complete electronics packaging which controls the Mars roving vehicle is described in order to provide a broad overview of the systems that are part of that package. Some software debugging tools are also discussed. Particular emphasis is given to those systems that are controlled by the microprocessor. These include the laser mast, the telemetry system, the command link prime interface board, and the prime software.
To identify structural requirements and technology needs of the geostationary platforms, two design alternatives are analyzed with respect to utilities accommodation, interface, and strength and stiffness requirements. Alternative 1 is comprised of 6,800 kg (15,000 lb) platform modules, each delivered to low earth orbit (LEO) with an attached orbit transfer vehicle (OTV) in a single Shuttle flight, deployed, and transferred to a geostationary constellation of platforms. Alternative 2 is made up of 16,800 kg (37,000 lb) platform modules, each delivered to LEO, deployed, mated to a 2 stage OTV (delivered to LEO in two additional Shuttle flights) and transferred to GEO for docking with other modules to form a single large platform. The top candidates for deployable structure elements are the expandable truss beam and the Astromast. The beam, fabricated of graphite/epoxy with a designed coefficient of thermal expansion of zero, provides good packaging density and high strength and stiffness per unit weight. The Astromast provides the best packaging density, but lacks the strength and stiffness required for the platform payloads. A graphite/epoxy mast with the Astromast packaging density is an advancement in technology needed for optimizing the geostationary platform designs.
A review of the NASA supplied mission scenarios for the communications, radiometry and radio astronomy missions. Led to specific hoop/column antenna configurations for each mission. The mission configurations were then evaluated to identify specific technology items requiring further development. The compilation of these technology drivers resulted in a specification of an artificial or point design, the design element around which all design and performance estimates for the rest of the program were made. Mechanisms for deployment and stowing are examined including the cable driven mast and the latch pulley roller. Methods for determining the thermal expansion of candidate materials, the development of the cables, cords, and hinged joints, and surface adjustment are considered.
Equations are developed for vibration of general lattice structures that have repetitive geometry. The method of solution is an extension of a previous paper for buckling of similar structures. The theory is based on representing each member of the structure with the exact dynamic stiffness matrix and taking advantage of the repetitive geometry to obtain an eigenvalue problem involving the degrees-of-freedom at a single node in the lattice. Results are given for shell-and beam-like lattice structures and for rings stiffened with tension cables and a central mast. The variation of frequency with external loading and the effect of local member vibration on overall modes is shown.
The atmospheric emissions photometric imaging experiment (AEPI) to be flown on Spacelab 1 is designed to study faint natural and artificial atmospheric emission phenomena. Optical emissions are imaged in the region 2150 A to 7320 A using a television system consisting of two optical channels, one wide-angle and one telephoto. The detection system is an image-enhanced SEC vidicon. A third optical channel images onto the photocathode of a microchannel plate photomultiplier tube that has 100 discrete anodes. Photons are counted for each discrete anode, providing a direct measure of the luminosity of an object viewed by the TV telephoto lens, albeit with low spatial resolution. The AEPI detector is mounted on a two-axis gimbal comprised of a Modified Apollo Telescope Mount Star Tracker (MAST), which provides experiment pointing over a 40-deg x 80-deg range, exclusive of restrictions due to the proximity of other experiments. The pointing stability is 1 arcmin with respect to the spacecraft coordinate system for an exposure of 1 second. The tracking capability is 3.5 deg/s with a stability of 1 arcmin. The detector and pointing system are located on the Spacelab pallet. The experiment is controlled by stored programs resident in the Dedicated Experiment Processor located in the Spacelab module.
Unfoldable-membrane-reflector concept leads to mobile photovoltaic generators. Hinged containers swing open for deployment, and counterbalance beam swings into position. Folded reflector membranes are unfolded as deployment mast is extended, until stretched out flat.
It appears practical to challenge the Orbiter DAP with a large, attached structure. The definition of this capability is a fundamental step in the development of nearly all large space systems currently under consideration. Experiment features may be incorporated that apply to control systems for large space systems such as modal damping devices. In a relatively simple deployable structure, the correlation of flight test results with ground test and analysis should provide a basis for extrapolation to more complex structures. Initial experiment concepts will provide a starting point for the examination of antenna feed mast requirements with the objective changing the design to produce a representative test article. Correlation of construction operations with ground simulations will provide for better task and time-line definition. EVA needs to be a direct benefit to the conduct of the experiment. Early consideration of safety issues is a precaution against defining an unacceptable experiment concept. Integration of many objectives seems feasible and is generally perceived as the only way to justify a relatively expensive experiment.
The simulated zero-gravity ground testing of the flexible fold-up solar array consisting of eighty-four full-size panels (.368 m x .4 m each) is addressed. Automatic, hands-off extension, retraction, and lockup operations are included. Three methods of ground testing were investigated: (1) vertical testing; (2) horizontal testing, using an overhead water trough to support the panels; and (3) horizontal testing, using an overhead track in conjunction with a counterweight system to support the panels. Method 3 was selected as baseline. The wing/assembly vertical support structure, the five-tier overhead track, and the mast-element support track comprise the test structure. The flexible solar array wing assembly was successfully extended and retracted numerous times under simulated zero-gravity conditions.
A deployable mast concept which meets the weight, size and stability requirements for a feed support structure for offset antennas up to 100 meters in diameter is discussed. A triangulated truss configuration, the use of tapered tubes which exhibit a high strength-to-weight ratio, and low CTE graphite-epoxy material are seen to provide an efficient, lightweight and stable truss suitable for an antenna feed support. A low stowage ratio of 30:1 is achieved through a unique preloaded hinge located at the center of each longeron and an autonomous deployment cage with a drive mechanism. Initial analysis and proof of concept hardware validated the basic mechanism and design assumptions and provided a basis for further investigation. The concept can readily accept variations in member size and thus lends itself to optimization for other potential uses where a stiff, lightweight deployable truss is needed.
The technology development needed to support the satellite system conceptual design is amplified. Activities connected to RF, control, and structure of the MSAT UHF antenna are addressed. The antenna RF activities are considered, and two studies related to the reflector surface tolerance and the microstrip feed technology are described. The software and hardware needed to implement the attitude control conceptual design are reviewed. Antenna structure development for the reflector and the deployable supporting mast is discussed. UHF feed located electronics are addressed.