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The probe bus and orbiter subsystems are defined, and tradeoffs analyzed. Subsystems discussed include: communications, electric power, data handling, attitude determination and control, propulsion, thermal control, structure and mechanisms, NASA/ESRO orbiter interface, mission operation, and flight support.
The design is described of the Venus probe windows, which are required to measure solar flux, infrared flux, aureole, and cloud particles. Window heating and structural materials for the probe window assemblies are discussed along with the magnetometer. The command lists for science, power and communication requirements, telemetry sign characteristics, mission profile summary, mass properties of payloads, and failure modes are presented.
The procedures for predicting the aeroheating environment of Venus entry probes are outlined. After some consideration, a number of assumptions were adopted in order to make the prediction techniques tractable. Among these assumptions are thermochemical equilibrium, uncoupled radiative and convective processes, and uncoupled ablation products effects. The single strip method of integral relations, appropriately constrained, is shown to provide adequate inviscid results as a basis for heating calculations on blunt configurations. Techniques for prediction of the laminar, transitional, and turbulent convective environment are outlined and shown to agree with data. The prediction of radiative heating in C, N, and O gas mixtures is discussed and a practical scheme adopted. A comparison with LRC calculations is made.
An analysis of low cost management approaches for the development of the Earth Observatory Satellite (EOS) is presented. The factors of the program which tend to increase costs are identified. The NASA/Industry interface is stressed to show how the interface can be improved to produce reduced program costs. Techniques and examples of cost reduction which can be applied to the EOS program are tabulated. Specific recommendations for actions to be taken to reduce costs in prescribed areas are submitted.
An analysis was conducted to determine the compatibility of the Earth Observatory Satellite (EOS) with the space shuttle. The mechanical interfaces and provisions required for a launch or retrieval of the EOS by the space shuttle are summarized. The space shuttle flight support equipment required for the operation is defined. Diagrams of the space shuttle in various configurations are provised to show the mission capability with the EOS. The subjects considered are as follows: (1) structural and mechanical interfaces, (2) spacecraft retention and deployment, (3) spacecraft retrieval, (4) electrical interfaces, (5) payload shuttle operations, (6) shuttle mode cost analysis, (7) shuttle orbit trades, and (8) safety considerations.
A study was conducted to determine the recommended orbit for the Earth Observatory Satellite (EOS) Land Resources Mission. It was determined that a promising sun synchronous orbit is 366 nautical miles when using an instrument with a 100 nautical mile swath width. The orbit has a 17 day repeat cycle and a 14 nautical mile swath overlap. Payloads were developed for each mission, EOS A through F. For each mission, the lowest cost booster that was capable of lifting the payload to the EOS orbit was selected. The launch vehicles selected for the missions are identified on the basis of tradeoff studies and recommendations. The reliability aspects of the launch vehicles are analyzed.
The key issues in the Earth Observatory Satellite (EOS) program which are subject to configuration study and tradeoff are identified. The issue of a combined operational and research and development program is considered. It is stated that cost and spacecraft weight are the key design variables and design options are proposed in terms of these parameters. A cost analysis of the EOS program is provided. Diagrams of the satellite configuration and subsystem components are included.
The results of structural studies of the Earth Observatory Satellite (EOS) which define the member sizes to meet the vehicle design requirements are presented. The most significant requirements in sizing the members are the stiffness required to meet the launch vehicle design frequencies both in the late al and in the longitudinal directions. The selected configurations, both baseline and preferred, for the Delta and Titan launch vehicles were evaluated for stiffness requirements. The structural idealization used to estimate the stiffness of each structural arrangement, was based on an evaluation of primary loads paths, effectivity of structural members, and estimated sizes for the preferred configurations. The study included an evaluation of the following structural materials: (1) aluminum alloys, (2) titanium alloys, (3) beryllium, (4) beryllium/aluminum alloy, and (5) composite materials.
A management approach for the Earth Observatory Satellite (EOS) which will meet the challenge of a constrained cost environment is presented. Areas of consideration are contracting techniques, test philosophy, reliability and quality assurance requirements, commonality options, and documentation and control requirements. The various functional areas which were examined for cost reduction possibilities are identified. The recommended management approach is developed to show the primary and alternative methods.
The impacts of achieving compatibility of the Earth Observatory Satellite (EOS) with the space shuttle and the potential benefits of space shuttle utilization are discussed. Mission requirements and mission suitability, including the effects of multiple spacecraft missions, are addressed for the full spectrum of the missions. Design impact is assessed primarily against Mission B, but unique requirements reflected by Mission A, B, and C are addressed. The preliminary results indicated that the resupply mission had the most pronounced impact on spacecraft design and cost. Program costs are developed for the design changes necessary to achieve EOS-B compatibility with Space Shuttle operations. Non-recurring and recurring unit costs are determined, including development, test, ground support and logistics, and integration efforts. Mission suitability is addressed in terms of performance, volume, and center of gravity compatibility with both space shuttle and conventional launch vehicle capabilities.
The modifications required of the Pioneer F/G spacecraft design for it to deliver an atmospheric entry probe to the planets Saturn and Uranus are investigated. It is concluded that it is feasible to conduct such a mission within the constraints and interfaces defined. The spacecraft required to perform the mission is derived from the Pioneer F/G design, and the modifications required are generally routinely conceived and executed. The entry probe is necessarily a new design, although it draws on the technology of past, present, and imminent programs of planetary atmospheric investigations.
Technical progress in each of the disciplinary research areas affecting the design of supersonic cruise aircraft is discussed. The NASA AST/SCAR Program supported the integration of these technical advances into supersonic cruise aircraft configuration concepts. While the baseline concepts reflect differing design philosophy, all reflect a level of economic performance considerably above the current foreign aircraft as well as the former U.S. SST. Range-payload characteristics of the study configurating show significant improvement, while meeting environmental goals such as takeoff and landing noise and upper atmospheric pollution.
The merits, uses, and benefits of the Manned Orbital Space Station to mankind are examined. Spacecraft and space station configurations are shown.
Variable cycle engines were identified, based on the mixed-flow low-bypass-ratio augmented turbofan cycle, which has shown excellent range capability in the AST airplane. The best mixed-flow augmented turbofan engine was selected based on range in the AST Baseline Airplane. Selected variable cycle engine features were added to this best conventional baseline engine, and the Dual-Cycle VCE and Double-Bypass VCE were defined. The conventional mixed-flow turbofan and the Double-Bypass VCE were on the subjects of engine preliminary design studies to determine mechanical feasibility, confirm weight and dimensional estimates, and identify the necessary technology considered not yet available. Critical engine components were studied and incorporated into the variable cycle engine design.
Results pertinent to predicting the performance of convolutionally encoded binary phase-shift keyed communication links were presented. The details of the development are provided in four sections. These sections are concerned with developing the bit error probability performance degradations due to PN despreading by a time-shared delay locked loop, the Costas demodulation process, symbol synchronization effects and cycle slipping phenomena in the Costas loop. In addition, Costas cycle slipping probabilities are studied as functions of Doppler count time and signal-to-noise conditions. The effect of cycle slipping in the symbol synchronizer is also studied as a function of channel Doppler and other frequency uncertainties.
Some of the more important methods are discussed that have been used or proposed for aircraft parameter identification. The methods are classified into two groups: Equation error or regression estimates and Bayesian estimates and their derivatives that are based on probabilistic concepts. In both of these two groups the cost function can be optimized either globally over the entire time span of the transient, or sequentially, leading to the formulation of optimum filters. Identifiability problems and the validation of the estimates are briefly outlined, and applications to lifting rotors are discussed.
For abstract, see N77-11084.
Various types of measurements were studied for estimating the orbit and/or attitude of an Earth Observation Satellite. An investigation was made into the use of known ground targets in the earth sensor imagery, in combination with onboard star sightings and/or range and range rate measurements by ground tracking stations or tracking satellites (TDRSS), to estimate satellite attitude, orbital ephemeris, and gyro bias drift. Generalized measurement equations were derived for star measurements with a particular type of star tracker, and for landmark measurements with a multispectral scanner being proposed for an advanced Earth Observation Satellite. The use of infra-red horizon measurements to estimate the attitude and gyro bias drift of a geosynchronous satellite was explored.