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One of the early flights of NASA's space shuttle will be allocated to the Langley Research Center as an orbiting research laboratory, and a study team of scientists in the Shuttle Experiments Office has the responsibility for selecting the set of scientific experiments which will be performed on the flights. Working closely with the team, the authors developed a procedure which leads to an `optimal' resolution of the group decision problem. The approach taken can be characterized as a way of organizing the expert judgments of the team members to establish a measure of group preference over the set of experimental packages, and a particular feature of the procedure is the manner in which it structured a hierarchical taxonomy of the group task and talent to be employed in combination with the assignment of `influence weights' to achieve a maximal utilization of the expertise of the study team. Though the information demands are somewhat high, the approach was fully accepted by the study team as a rational way to make the selection of an experimental package and considerably preferred to an essentially political mode of resolution.

A guidance method for the space shuttle's transition from hypersonic entry to subsonic cruising flight is presented. The method evolves from a numerical trajectory optimization technique in which kinetic energy and total energy (per unit weight) replace velocity and time in the dynamic equations. This allows the open end-time problem to be transformed to one of fixed terminal energy. In its ultimate form, E-Guidance obtains energy balance (including dynamic-pressure-rate damping) and path length control by angle-of-attack modulation and cross-range control by roll angle modulation. The guidance functions also form the basis for a pilot display of instantaneous maneuver limits and destination. Numerical results illustrate the E-Guidance concept and the optimal trajectories on which it is based.

A preliminary design of a manipulator system, applicable to a Free-Flying Teleoperator Spececraft operating in conjunction with the Shuttle or Tug, is presented. The manipulator arm incorporates two 4-ft segments to the wrist with actuators located at the shoulder, elbow, and wrist. The wrist provides three degrees-of-freedom through pitch, yaw and continuous roll joints. An interchangeable end effector provides multiple task performance and satellite worksite versatility. A tip force of 10 1bs and a torque of 15 ft-1bs is provided. Man-in-the-loop simulations, using both unilateral and bilateral control techniques, were conducted. Based upon the simulation, a new, but relatively simple, control technique was proposed for the manipulator system.

Early in the 1980's NASA plans to orbit a Large Space Telescope using Space Shuttle as the launch platform. This orbiting telescope will permit scientists to observe the universe with a clarity never before attained through ground systems inhibited by the Earth's atmosphere. Current design goals indicate that the Telescope is to be designed for a fifteen year life span with the capability for on-orbit maintenance and, if required, earth return for servicing and refurbishment. Thus, NASA has decided to design a spacecraft for on-orbit EV manned maintenance thereby anticipating significant overall program cost savings. The design for orbital EV maintenance reaches across nearly every subsystem providing an excellent opportunity for complete human factors (crew systems) participation from system concept through design, integration, test, and ultimate operation. This shift in approach to on-orbit maintenance has resulted in major challenges such as designing for suited astronaut access to all subsystem equipment elements, minimization of contamination, handling of extremely sensitive instruments and their transfer through and within minimum volumes, development of translation techniques, incorporation of built-in crew work aids recognizing weight/volume limits and accomplishing the maintenance cycle within timelines imposed by mission, suit, and changeout constraints. These and other considerations will be discussed, alternative solutions presented, and recommendations made with particular reference to those factors which must be undertaken early to assure a timely and fundamentally sound crew systems program.

Substantial cost savings and performance improvement will result by the use of Extravehicular Activity (EVA) to supplement or replace automation. Taking an all-pallet version of Langley Research Center's Advanced Technology Laboratory payload as an example, $54.5 million should be saved by EVA over automation, considering deployment and stowing only. Additional savings should accrue when reduced-reliability equipment (where permitted) is substituted for high reliability equipment and EVA is used for repairs. More comprehensively, launch and operation costs could also be reduced by elimination of the need to return to the ground for repairs; and production spending might be reduced when an entire vehicle was saved by manned EVA repair not feasible via automation. Potential disadvantages include increased cost due to development and manufacture of EVA equipment, payload provisions to enable EVA interfaces, training, orbiter modification, and prevention of EVA-caused contamination. Possible applications to the Space Shuttle missions are discussed.

Extravehicular activity (EVA) is a baselined Shuttle Orbiter capability. It can provide cost-effective servicing operations for payloads, on-orbit repair, and techniques for meeting various on-orbit contingencies. The basic EVA system planned consists of Orbiter-integral provisions, carry-on life support provisions, and EVA support equipment provisions. The Orbiter-integral provisions include the internal airlock, the EVA Egress module, and the docking module, each of which provides an egress path. The carry-on life support provisions will be two Extravehicular Mobility Units (EMU), composed of an Astronaut Life Support Assembly (ALSA) and a Space Suit Assembly (SSA). The support equipment includes tethers, special tools, translation aids, and workstations. The Shuttle Orbiter EVA requirements and possible tasks, EVA crew rescue equipment and techniques, crew support, EVA training, and EVA timeline activities are described.

SummaryThe equations of motion of a vehicle entering the earth’s atmosphere (space shuttle) are expanded to give the incidence oscillation equation. The stability conditions for small free oscillations are deduced. These stability conditions apply equally to short-period and long-period oscillations. A special case of critical damping is investigated and more explicit stability conditions are derived for that case.

A heat-pipe-cooled Space Shuttle orbiter wing leading edge was compared and evaluated against three alternate leading edge candidates: a refurbishable ablative design, and two other reusable versions employing coated columbium and carbon-carbon high-temperature segments. Each candidate concept was shown feasible in the Phase B environment. The reusable versions were all found to cost nearly the same and substantially less than the nonreusable ablative version. However, since current Shuttle trajectories produce higher heating rates than those used in this study, the carbon-carbon and heat pipe concepts are now the prime candidates for Shuttle leading edge thermal protection.

The accommodation requirements for a representative spectrum of automated spacecraft are analyzed and compared to the services provided by the Space Shuttle and Space Tug. The purpose and engineering characteristics of a number of possible items of general purpose Shuttle-payload interface equipment are discussed. The benefits provided to automated payloads by the Space Transportation System are summarized.

An analysis of the continuum inviscid and boundary-layer flowfields around various configurations representative of an early version of a space shuttle orbiter re-entering the atmosphere at 40° and 60° angle of attack is presented. The inviscid flow fields were predicted by 2-D time-dependent and 3-D steady finite difference/artificial viscosity techniques, together with a 2-D shock layer analysis method. Boundary-layer flow and heating rates were computed with an integral matrix technique. Configurations which were considered in the analysis included: 1) 2-D flat plate representations of the flat underside of the orbiter fuselage, 2) 2-D orbiter fuselage cross sections, 3) 2-D orbiter wing airfoils, and 4) a 3-D straight-wing orbiter configuration. Theoretical results correlated satisfactorily with pressure and heat transfer data on flat plates and wings (including shock intersection regions).

The world-wide Omega navigation system, used in conjunction with locally-derived differential Omega updates, promises to provide the position-fixing accuracy and reliability required 1.) by commercial aircraft for area navigation in even heavily-trafficked terminal locales, 2.) by military aircraft for the successful execution of the vast majority of tactical missions, and 3.) by space shuttle and similar vehicles for re-entry navigation during the post-blackout portion of flight. A mechanization of the differential Omega concept is presented which provides maximum flexibility, allowing users to determine their positions with circular (rho-rho), hyperbolic, or pure differential mechanizations and yet does not require the differential ground subsystem to contain an expensive, highly-accurate oscillator which is normally associated with rho-rho capabilities. The hardware and software implementations are outlined and a rationale for the particular navigation and updating mechanization selected is presented. It is shown that while skywave propagation variations and local oscillator off-sets are not separated by the ground subsystem, this ambiguity need not introduce errors to user aircraft. In fact, users operating in the circular mode can generally achieve greater accuracy than for a purely hyperbolic mechanization in which all clock errors were eliminated. The implementation described is incorporated as a selectable mode of operation in the Omega subsystems to be delivered by Bendix Navigation and Control to the FAA for its differential Omega evaluation program.

This paper discusses the results of employing a first-order optimization algorithm to the problem of minimizing the weight of the re-entry thermal protection system (TPS) for the space shuttle. Mathematical models of two types of thermal protection systems are derived, a metallic TPS and a reusable surface insulation (RSI) TPS. Optimal entries were generated using maximum orbiter nose temperature as a parameter. Thermal protection system weights were computed for both fixed and variable angles of attack using three-dimensional entry trajectories. Results indicated that variable angle-of-attack entries require less thermal protection system weight than entries at a constant angle of attack (35°) for both systems considered.

The work discussed in this paper is directed toward possible environmental effects of high-altitude craft such as the SST and the space shuttle vehicle. The model includes the chemistry of the O-H-N system and vertical eddy transport. When the predicted NOX emission from an SST fleet projected for 1990 is introduced, it is found that the reduction in ozone column density is less than 15%. Furthermore, flight altitudes near 16 km lead to ozone reductions which are much less than are associated with flight altitudes near 20 km. The effect of projected space shuttle operations on mesospheric composition is negligible.

Some of the early results reported by several investigators in the solar physics discipline are described, as well as a subjective response to long-term zero-gravity exposure in several areas related to medical studies. Many of the roles in which man has participated in research aboard Skylab will again be important in future space activities such as the Space Shuttle. Examples are pointing, alignment, transient identification and response, and assurance of data quality. These tasks can be extended to include a greater role in data management and interpretation, so that some selectivity or processing, or both, might be appropriate onboard the spacecraft, prior to transmission of data to the ground.

The fracture properties of an ultralow-expansion glass intended for use in windows for the Space Shuttle were characterized by strength and fracture-mechanics techniques to provide reliable design data. Proof-test diagrams for predicting minimum times-to-failure under specified service loads were developed from measurements of subcritical crack growth in water and air. Failure predictions were confirmed from strength measurements in water. In vacuum (

The determination of optimal rocket control profiles and the guidance* procedures used to approximate these profiles have received much attention during the last decade. Simple guidance procedures have sufficed for the Apollo flights. However, for the space shuttle which will have a low-thrust, orbital maneuvering system a more sophisticated guidance is required. This paper develops a guidance applicable to low-thrust orbital maneuvers and compares numerical results with those obtained using lengthy gradient and accelerated gradient methods.

Where high-level technology is concerned, questions of cost-benefits and funding now often rival the once dominant hardware and engineering problems as major program challenges. Space shuttle proponents point to significant savings expected over the comparatively inefficient Saturn/Apollo launch and recovery system and indicate that a wide cross section of scientific interests will profit from direct access to extraterrestrial experimentation. Even engineering work on unmanned flights is geared toward squeezing a better power/weight ratio out of solar cells, storage batteries, and power control electronics. As a direct result of the Mideast war, budgeting for U.S. military R&D must now confront the immediate challenges of new Soviet ECM and their SA-6 surface-to-air missile as opposed to projected and perhaps speculative benefits claimed for the billion dollar Trident submarine program.

References 1 "Apollo 16 Brings Us Visions from Space," National Geographic, Vol. 142, No. 6, Dec. 1972, pp. 856-865. 2 Baldwin, H. L., "Orbit Display Using the SC-4020," Kept. 63-0459, 1963, General Dynamics/Astronautics, San Diego, Calif. 3 Hitzl, D. L., "Computer Graphics," Internal Note, Oct. 1969, Lockheed Palo Alto Research Lab., Palo Alto, Calif. 4 Hitzl, D. L. and Shayer, S., "Computer Illustrations," Internal Note, Oct. 1969, Lockheed Palo Alto Research Lab., Palo Alto, Calif. 5 Hitzl, D. L., "The Relative Motion of Commensurate Earth Satellites," Internal Note, Aug. 1971, Lockheed Palo Alto Research Lab., Palo Alto, Calif. b Madigan. J. T.< "Orbit Display Using the SC-4020 Subroutine MPLOT," Lockheed Missiles and Space Co., Sunnyvale, Calif., 1963.

Statistical estimates of wind shear in the planetary boundary layer are important in the design of V/STOL aircraft, and for the design of the Space Shuttle. The data analyzed in this study consist of eleven sets of longitudinal turbulent velocity fluctuation time histories digitized at 0.2 sec intervals with approximately 18,000 data points per time history. The longitudinal velocity fluctuations were calculated with horizontal wind and direction data collected at the 18-, 30-, 60-, 90-, 120-, and 150-m levels. The data obtained confirm the result that Eulerian time spectra transformed to wave-number spectra with Taylor's frozen eddy hypothesis possess inertial-like behavior at wave-numbers well out of the inertial subrange.

An optimal Space Shuttle ascent-flyback trajectory shaping capability is presented which is based on the accelerated gradient parameter optimization technique. A typical atmospheric flight branched optimization problem is analyzed which required the determination of 31 parameters. This parameter set includes the description of the vehicle attitude control angles for three branches of Shuttle flight: first stage ascent, second stage ascent, and first stage flyback. The important in-flight inequality constraints required to maintain the integrity of the vehicle are considered. Results indicate that for a launch into a 55 deg inclined ellipse, a 13% increase in payload can be realized by using optimal control in the first-stage ascent rather than the conventional gravity-turn steering.

Abort is defined as a premature or abnormal termination of a mission as a result of either a system or human failure. The abort guidance during the ascent phase, among all flight phases of the shuttle vehicle, is most critical; because it involves guiding a heavily loaded shuttle vehicle that is passing through severe environmental conditions. This paper describes a predictive guidance algorithm for the abort mode initiated during the orbiter ascent flight phase. Contrary to the concept of the minimization1 of fuel consumption during a nominal ascent flight phase, the Space Shuttle orbiter is required to deplete its fuel tank in order to be separated from the vehicle during the abort mode. The guidance objective during the abort mode is to attain a proper terminal energy state in relation to the landing site, consistent with the aerodynamic heating and loading constraints. In the absence of a well-defined payoff function for optimization during this mode, the abort guidance algorithm presented in this paper is formulated to minimize aerodynamic forces. Minimization of aerodynamic forces was considered because of the possibility that the failure that caused the initiation of the abort mode may have weakened the orbiter structure. Immediately after the abort mode initiation, consistent with the philosophy of the minimization of the aerodynamic forces, the orbiter is forced to fly a lifting trajectory to attain higher altitude, thereby reaching a low dynamic pressure environment. The predictive guidance algorithm uses the knowledge of the current orbiter state and the desired terminal state in relation to the landing site, and determines the time sequence of the powered maneuver and the orientation of the thrust vector throughout the abort mode. For example, the magnitude of the down-range component of the acceleration, the time at which to begin to return toward the desired landing site, and the radial acceleration necessary to attain the desired terminal altitude (while remaining in the low dynamic pressure environment) are computed. The key formulations in the predictive guidance algorithm are analytical expressions for the terminal state that the vehicle will reach for a given sequence of thrust orientations, and the associated Jacobian matrix. Therefore, these have been carefully developed in the paper. In the terminal phase of the abort mode, the predictive guidance reduces to a velocity to be gained scheme which has the desirable effect of reducing the deviations in the end-state prediction that results from variations and uncertainties in the vehicle characteristics. Detailed results for representative abort cases and estimates of the computer requirements are also included in the paper.

If R. J. Mitchell were still alive, then it is very probable that this creator of the Spitfire, the Schneider Trophy seaplanes, and other classic aircraft would be attracted by the challenge of space flight. This is more than ever likely at this precise point in time, when the technology of the aeroplane is being joined with that of the ballistic rocket, in readiness to take the next great step forward in astronautics with the new vehicles which have come to be called “space shuttles”.It was with these thoughts in mind that I chose the subject matter of this Memorial Lecture, when the Southampton Branch honoured me with the invitation to deliver it.

An optimal atmospheric flight branched trajectory-shaping capability is presented based on the Davidon-Fletcher-Powell variable metric parameter optimization technique. Gradient information is generated using finite difference methods. A typical atmospheric flight branched optimization problem is analyzed which requires the determination of 31 parameters. This parameter set includes the three-dimensional description of vehicle attitude control angles for three branches of flight: first-stage ascent, second-stage ascent, and first- stage flyback. The important inflight inequality contraints required to maintain the integrity of the vehicles are considered. Some of the numerical methods employed are discussed, along with several new auxiliary techniques developed to improve the compatibility of the numerical gradient and iterator.

Precision distance measuring equipment (DME) aiding onboard inertial navigation can be used to meet the stringent landing navigation requirements of the Space Shuttle. The required accuracy is achieved by using a ten-state-variable Kalman filter to process the measured range and delta-range data from the DME. The DME transponders are required to meet the accuracy requirements: two are deployed under the final approach path and one is deployed to the side. Failure considerations require deployment of additional redundant transponders. Initialization of the landing navigation after hypersonic entry is accomplished using the transponders at the airport.

The new space transportation system based on a reusable space shuttle, currently under development, will substantially change our traditional methods of space research that have been based on a ballistic‐missile technology and its reliance on expensive expendable components. The shuttle will not become operational until 1979, but the time is ripe now for planning the space‐exploration and space‐science programs for the period 1978–85.

Integrated electronics for NASA space shuttle program, discussing costs, configurations and technologies application

The space shuttle will re-enter the Earth's atmosphere at high angle of attack, in order to minimize total heat load, and will perform a transition to low angle of attack for increased trajectory control. The timing and shaping of this angle transition are affected by specific-control-moment boundaries, attitude instability, angle of attack, limits on various flight parameters, and the establishment of a trimmed condition for flight to the landing site. Transitions to altitudes between 45,000 and 150,000 ft are investigated here, with particular attention given to suitable terminal conditions and to stability and control boundaries. Angle of attack is the control variable; hence, the short-period dynamics are not modeled. Two degree-of-freedom trajectories are optimized using a steepest descent algorithm with automatic step-sizing and open end-time. It is found that trim-glide flight-path angle y becomes very steep as altitude increases and that trajectories to such flight conditions are difficult to achieve. High-altitude equilibrium conditions (y = 0) at zero y are readily obtained, and they provide suitable way points for a continued glide to low altitude. Dynamic pressure and load factor peaks associated with both types of transition are not excessive. Stability and control boundaries have significant effects on the trajectories, although peak load, dynamic pressure and load factor are not materially altered. Gliding trajectories from high to low altitude are found to be sensitive to phugoid mode excitation, and the dynamic pressure peak which occurs prior to the terminal flight condition may be excessive.

This paper studies the feasibility of the use of an atmospheric rendezvous concept to increase the efficiency and flexibility of space transportation systems. In this concept the second stage of a recoverable orbital launch vehicle or hypersonic transport would be built without wings, landing gear, or subsonic flight propulsion, and would be received in an atmospheric rendezvous by a carrier vehicle at the terminal point of flight for subsequent ferry to a landing site. Significant possibilities for weight improvement are shown and the attractiveness of a subsonic form of atmospheric rendezvous in either a towing or docking mode is illustrated.

Unmanned and manned interorbital shuttle systems for satellites placement and repair and planetary spacecraft insertion, emphasizing cost analysis

Space based reusable manned/unmanned tug, discussing potential missions, system requirements and auxiliary hydrogen oxygen propulsion system

3 Hanna, G. L., Troiano, A. R., and Steigerwald, E. A., "A Mechanism for Embrittlement of High Strength Steels by Aqueous Environments," Transactions American Society for Metals, Vol. 57, 1964, pp. 658-671. 4 Dull, D. L. and Raymond, L., "A Method of Evaluating Relative Susceptibility of Bolting Materials to Stress Corrosion Cracking," presented at 1972 Westec Conference, March 1972; also Corrosion, to be published. 5 Phillips, A., Kerlins, V., and Whiteson, B. V., "Electron Fractographic Handbook," AFML Rept. ML-TDR-64-416, Aug. 1968, Air Force Materials Lab., Wright-Patterson Air Force Base, Ohio. 6 Fidelle, J. P., Legrand, J., and Couderc, C., "A Fractographic Study of Hydrogen Gas Embrittlement in Steels," Transactions of the American Institute of Mining, Metallurgical Engineers, 1972, to be published. 7 Tiner, N. A. and Filpin, C. B., "Microprocesses in Stress Corrosion of Martensitic Steels, Corrosion, Vol. 22, 1966, pp. 271279. 8 Brown, B. F., "A New Stress Corrosion Cracking Test for High Strength Alloys," Materials Research and Standards, Vol. 16, March 1966, pp. 129-133. 9 Brown, B. F., "The Application of Fracture Mechanics to Stress-Corrosion Cracking," Metallurgical Review, Vol. 13, 1968, pp. 171-183. 10 Mulherin, J. H., "Stress Corrosion Susceptibility of High Strength Steel in Relation to Fracture Toughness," Transactions of the ASME: Journal of Basic Engineering, Vol. 88, 1966, pp. 772782. 11 Tentative Method of Test for Plane-Strain Fracture Toughness of Metallic Materials, E399-70T, Book of Standards, American Society for Testing and Materials, Philadelphia, Pt. 31, July 1971, pp.911-1927. Freedman, A. J., "Development of an Accelerated StressCorrosion Test for Ferrous and Nickel Alloys," Finary Summary Rept., NOR 68-58, April 1968, Northrop Corp., Hawthorne, Calif.

Reusable air breathing launch vehicle for earth orbit shuttle, comparing performance, costs and operation with rocket powered systems

Low-cost transportation to and from Earth orbit with systems characterized by reusability, flexibility, and "airline-type" operations is a prime goal of the space program. The airbreathing HTOL launch vehicle has inherent features which make it a candidate for such a system but, unfortunately, the lack of a developed powerplant implies a procurement cycle length on the order of a decade. The purpose of this paper is to assess the status of the airbreathing launch vehicle in the light of recent achievements in propulsion and configuration development. It is concluded that all maj or questions of feasibility, particularly in the area of hypersonic propulsion, have been answered favorably by the investigations of several small research engines recently completed. The remaining key problems are discussed together with some promising approaches toward their solution. Finally, important steps required to develop an operational air-breathing system by the early eighties are discussed.

Reusable nuclear shuttle /RNS/ for space program providing low cost transportation beyond earth orbit

Theoretical structural models are developed for point supported panels typical of proposed space shuttle construction which are suitable for vibration and flutter analyses. Representative results for panel natural frequencies and flutter boundaries are presented. These are obtained using quasi-steady aerodynamics which ignore the aerodynamic boundary layer. It is shown for a proposed space shuttle design that the supports must be modeled as points rather than as continuous supports in order to obtain accurate results and that simply increasing the number of supports may not be beneficial as far as flutter is concerned.