ALBERT

Description: A KC-135A aircraft equipped with wing tip winglets was flight tested to demonstrate and validate the potential performance gain of the winglet concept as predicted from analytical and wind tunnel data. Flight data were obtained at cruise conditions for Mach numbers of 0.70, 0.75, and 0.80 at a nominal altitude of 36,000 ft. and winglet configurations of 15 deg cant/-4 deg incidence, 0 deg cant/-4 deg incidence, and baseline. For the Mach numbers tested the data show that the addition of winglets did not affect the lifting characteristics of the wing. However, both winglet configurations showed a drag reduction over the baseline configuration, with the best winglet configuration being the 15 deg cant/-4 deg incidence configuration. This drag reduction due to winglets also increased with increasing lift coefficient. It was also shown that a small difference exists between the 15 deg cant/-4 deg incidence flight and wind tunnel predicted data. This difference was attributed to the pillowing of the winglet skins in flight which would decrease the winglet performance.

Description: A joint NASA/USAF program was conducted to accomplish the following objectives: (1) evaluate the benefits that could be achieved from the application of winglets to KC-135 aircraft; and (2) determine the ability of wind tunnel tests and analytical analysis to predict winglet characteristics. The program included wind-tunnel development of a test winglet configuration; analytical predictions of the changes to the aircraft resulting from the application of the test winglet; and finally, flight tests of the developed configuration. Pressure distribution, loads, stability and control, buffet, fuel mileage, and flutter data were obtained to fulfill the objectives of the program.

Description: A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.

Description: A comprehensive set of measurements about the orbiter environment are provided by the plasma diagnostics package (PDP). Ion and electron particle densities, energies, and spatial distribution functions; ion mass for identification of particular molecular ion species; and magnetic fields, electric fields and electromagnetic waves over a broad frequency range are studied. Shuttle environmental measurements will be made both on the pallet and, by use of the remote manipulator system (RMS), the PDP will be maneuvered in and external to the bay area to continue environmental measurements and to carry on a joint plasma experiment with the Utah State University fast-pulsed electron generator. Results of orbiter environment EMI measurements and S-band field strengths as well as preliminary results from wake search operations indicating wake boundary identifiers are reported.

Description: The Earth radiation budget experiment (ERBE) software development approach is described. An iterative development approach was adopted which provides for three releases or versions of the processing system, each of increasing levels of complexity and solidity. The final release of the system will be used to process the flight data. The major phases for each iterative release consist of specifications developed in concert with the science team, preliminary design, subsystem reviews, coding, subsystem code walkthroughs, system testing, system documentation, and project status review.

Description: The correlation of outgassing to the stability of the damping properties of polymer materials to be used in spacecraft structures is discussed. A test series was devised to obtain basic information from off-the-shelf damping materials. The test results could be considered as a guideline toward the application of these materials. Eight materials were selected to form a representative cross section of those polymers having both ready availability as commercial damping materials and desirable properties. A table indicates the temperatures at which peak damping occurs at 1 Hz and the type of beam specimen used in the vacuum exposure tests. These materials as a group cover the temperature range of -85 C to 38 C.

Description: A molecular beam facility to simulate the space environment of a spacecraft at low orbit was designed with the intent of studying the effect on the properties of optical elements of oxygen atoms impacting at orbital velocity. The four-stage differentially pumped molecular beam facility includes a variety of oxygen atom beam sources which cover a wide range of velocities (1 km/sec to approximately 8 km/sec), in addition to the ultra-clean experimental environmental of an ultra-high vacuum chamber and an optical diagnostic set-up. The primary oxygen atom beam source used to obtain the 8 km/sec O atoms is an arc heated source. It consists of a modified commercially available plasma torch. The modifications include attachments which provide for a nozzle which is used to expand the atomic beam into the vacuum system, and exhaust channels to dispose of excess torch gas. The torch operates in the 'nontransferred' mode of operation, that is the electric arc is confined within the torch. A plasma is formed in helium by a dc arc. A small amount of O2 is injected downstream from the arc where it is thermally dissociated by the hot He into oxygen atoms. The high temperature and isentropic expansion give the oxygen atoms their velocity. Using seeded beam techniques, oxygen atom beams of approximately 3.5 and approximately 1.5 km/sec, respectively, are obtained.

Description: A preliminary assessment of the space shuttle contamination environment was made using data from the first two orbital flight tests, STS-1 and STS-2. Data sources consisted of crew observations during flight, postflight vehicle inspection, and the induced environment contamination monitor which was used on STS-2 and consists of 10 instruments. These instruments are used to measure gas phase contaminants, particle population, humidity, and molecular deposition in the orbiter payload bay during ascent and descent and particle population, molecular deposition, and gas cloud during orbital flight. Results of the measurements described are presented in summary form and indicate low molecular deposition rates for both pressurized and orbital flight.

Description: The structure of the European Space Agency (ESA) is described. The major test facilities used for ESA programs are described. Facility characteristics and special test methods are described.

Description: The training program and procedures developed and implemented at the space simulation laboratory at Martin Marietta Aerospace in Denver are discussed. The training of technicians and professionals as well as preparation for instructors is covered. Training manuals and their compilation are reported as applicable to the specific needs of the laboratory. The development of a space simulation course as part of the Martin Marietta Continuing Education Night School approaching space simulation from an academic viewpoint is presented. Finally, public relations tours of the facility as an informational/educational tool are discussed.

Description: A test involving a flow of hot hydrazine decomposition products at a rate of 13.6 g/s (0.03 lb/s) established the requirement to maintain pressure in an 11.9-m-(39 ft) diameter space chamber below 200 microns. The flow, 2/3 hydrogen and 1/3 nitrogen by volume, continued for several periods ranging from 3 to 15 min. The pressure requirement was necessary to minimize thermal effects of the gas on the test vehicle but was well beyond the capability of the existing facility pumps. Various methods of obtaining additional temporary jump capacity were considered. From these, the slugged-charcoal approach was selected as the quickest and least expensive method to implement.

Description: A thermal balance test (controlled flux intensity) on a simple black dummy spacecraft using IR lamps was performed and evaluated, the latter being aimed specifically at thermal mathematical model (TMM) verification. For reference purposes the model was also subjected to a solar simulation test (SST). The results show that the temperature distributions measured during IR testing for two different model attitudes under steady state conditions are reproducible with a TMM. The TMM test data correlation is not as accurate for IRT as for SST. Using the standard deviation of the temperature difference distribution (analysis minus test) the SST data correlation is better by a factor of 1.8 to 2.5. The lower figure applies to the measured and the higher to the computer-generated IR flux intensity distribution. Techniques of lamp power control are presented. A continuing work program is described which is aimed at quantifying the differences between solar simulation and infrared techniques for a model representing the thermal radiating surfaces of a large communications spacecraft.

Description: The Jet Propulsion Laboratory 25-foot space simulator with its 5.8-m (19-ft) diameter simulated solar beam provides an excellent facility for measuring the optical characteristics of parabolic solar concentrator panels and gores. The virtual source position and size were determined by using a single lamp of the 37 xenon 30-kW source array with only the center lens in the 19-channel optical mixer. This data was used to define the optical test geometry, and it allowed accurate measurement of focal length and surface deviations of the mirror under test. A flux distribution of a typical solar concentrator placed directly on the solar beam gives measurements of performance at the focal point of the parabolic surface.

Description: Several mathematical models, including a minimum integral square criterion problem, were used for the qualitative investigation of fuel optimal maneuvers for spacecraft with fixed thrusters. The solutions consist of intervals of "full thrust" and "coast" indicating that thrusters do not need to be designed as "throttleable" for fuel optimal performance. For the primary model considered, singular solutions occur only if the optimal solution is "pure translation". "Time optimal" singular solutions can be found which consist of intervals of "coast" and "full thrust". The shape of the optimal fuel consumption curve as a function of flight time was found to depend on whether or not the initial state is in the region admitting singular solutions. Comparisons of fuel optimal maneuvers in deep space with those relative to a point in circular orbit indicate that qualitative differences in the solutions can occur. Computation of fuel consumption for certain "pure translation" cases indicates that considerable savings in fuel can result from the fuel optimal maneuvers.

Description: A planar model of a space base and one module is considered. For this simplified system, a feedback controller which is compatible with the modular construction method is described. The systems dynamics are decomposed into two parts corresponding to base and module. The information structure of the problem is non-classical in that not all system information is supplied to each controller. The base controller is designed to accommodate structural changes that occur as the module is added and the module controller is designed to regulate its own states and follow commands from the base. Overall stability of the system is checked by Liapunov analysis and controller effectiveness is verified by computer simulation.

Description: A residual gas analyzer (RGA), a device for measuring the amounts and species of various gases present in a vacuum system is discussed. In a recent update of the RGA, it was shown that the use of microprocessors could revolutionize data acquisition and data reduction. This revolution is exemplified by the Inficon 1Q200 RGA which was selected to meet the needs of this update. The Inficon RGA and the Zilog microcomputer were interfaced in order the receive and format the digital data from the RGA. This automated approach is discussed in detail.

Description: By using infrared images obtained from GOES satellite, the digital count values of pixels representing blackbody temperatures of the cloud top, convective storms are observed throughout their life cycles. Clouds associated with a tornadic storm are compared with those without a tornadic storm to illustrate how the infrared and visible observations from a geosynchronous satellite can be used to study the differences in their life cycles.

Description: A space based orbital transfer vehicle (SBOTV) and ground based OTV's (GBOTV) are compared for debris protection, space based OTV maintenance provisions, flight performance, onorbit refueling, and launch and return operations. Debris protection has a severe impact on the SBOTV, while the penalty for the GBOTV is much less severe. A key technology issue is the protection capability of composite materials. Onorbit maintenance is critical for SBOTV. Reduction of losses during the various transfers is th maine problem with refueling a SBOTV. Zero-g propellant storage and transfer is an important technology area for SBOTV. A reusable shroud must be developed to return GBOTV's if a Shuttle derivative vehicle is used. The advantage of space basing lies in more efficient use of the launch vehicle. Since most of the mass going to LEO is OTV propellant, and the launches to deliver the SBOTV propellant are generally mass limited, substantially fewer launches are required to support the SBOTV.

Description: An environmental noise assessement of the initial launch of the Space Transportation System, STS-1 Columbia was conducted. The principal objective of the environmental noise assessment was to measure the noise generated during the initial launch of the space shuttle to ascertain the validity of the levels predicted in the 1979 environmental impact statement. In the 1979 study information obtained for expendable launch vehicles, Titan, Saturn and Atlas was used to predict the noise levels that would be generated by the simultaneous firing of the two solid rocket boosters and the three space shuttle main engines. Fifteen monitoring sites were established in accessable areas located from 4,953 to 23,640 meters from the launch pad. Precision sound level meters were used to capture the peak level during the launch. Data obtained was compared to the predicted levels and were also compared to the identified levels, standards and criteria established by the federal agencies with noise abatement and control responsibilities.

Description: Multilayer insulation blankets used for the attenuation of radiant heat transfer in spacecraft are addressed. Typically, blanket effectiveness is degraded by heat leaks in the joints between adjacent blankets and by heat leaks caused by the blanket fastener system. An approach to blanket design based upon modular sub-blankets with distributed seams and upon an associated fastener system that practically eliminates the through-the-blanket conductive path is described. Test results are discussed providing confirmation of the approach. The specific case of the thermal control system for the optical assembly of the Space Telescope is examined.

Description: Research efforts are reviewed on the space durability of materials, including radiation effects on polymer matrix composites and films, dimensional stability of polymer matrix composites and tension-stabilized cables, and thermal control coatings. Research to date has concentrated on establishing a fundamental understanding of space environmental effects on current graphite-reinforced composites and polymer systems, and development of analytical models to explain observed changes in mechanical, physical, and optical properties. As a result of these research efforts, new experimental facilities have been developed to simulate the space environment and measure the observed property changes. Chemical and microstructural analyses have also been performed to establish damage mechanisms and the limits for accelerated testing. The implications of these results on material selection and system performance are discussed and additional research needs and opportunities in the area of tougher resin/matrix and metal/matrix composites are identified.

Description: The paper describes the design of the solar array system (SAS) for the Solar Maximum Mission, the unique features of the SAS, and the results of its successful in-orbit operation. It is noted that the array was unique in that: (1) major weight concessions were made to produce a dynamically stiff array; (2) it was the first array designed to be compatible with the NASA Multimission Modular Spacecraft; (3) it is the first jettisonable solar array; and (4) it represents the first use of FEP-bonded overslides on a prime power array. It is concluded that the array performed as predicted with no evidence of the FEP causing any unusual array power degradations. In addition, the deployment and telemetry systems performed as designed.

Description: A broad overview is presented of current and near-term solar array technology that could be suitable for space use. Particular consideration is given to such advanced concepts as high power arrays, concentrator arrays, and ultrathin solar cell arrays. It is concluded that if such ambitious concepts as geosynchronous space platforms, orbital space stations, and alternate forms of propulsion are realized, the type of new technology described in this paper may find acceptance for space.

Description: The feasibility and performance parameters for beam microwave power supplies from a space station to nearby orbiting satellites are examined. A 5.8 GHz transmission frequency is found suitable for beaming 1-10 kW over a distance of 1-10 km. The antenna could have a 15 m diameter, a 64 kW output, provide uniform illumination, and have a retrodirective phase control system. A LEO to ground demonstration project is described, involving power levels of 0.0025 mW/sq cm and yielding 202 W at a 100 x 100 m rectenna.

Description: A Programmable Power Processor (P3) has been developed for application in future large space power systems. The P3 is capable of operation over a wide range of input voltage (26 to 375 Vdc) and output voltage (24 to 180 Vdc). The peak output power capability is 18 kW (180 V at 100 A). The output characteristics of the P3 can be programmed to any voltage and/or current level within the limits of the processor and may be controlled as a function of internal or external parameters. Seven breadboard P3s and one 'flight-type' engineering model P3 have been built and tested both individually and in electrical power systems. The programmable feature allows the P3 to be used in a variety of applications by changing the output characteristics. Test results, including efficiency at various input/output combinations, transient response, and output impedance, are presented.

Description: The group of techniques that as a class are referred to as synthetic battery cycling are described with reference to spacecraft battery systems. Synthetic battery cycling makes use of the capability of computer graphics to illustrate some of the basic characteristics of operation of individual electrodes within an operating electrochemical cell. It can also simulate the operation of an entire string of cells that are used as the energy storage subsystem of a power system.

Description: Rechargeable nickel-hydrogen systems are described that more closely resemble a fuel cell system than a traditional nickel-cadmium battery pack. This was stimulated by the currently emerging requirements related to large manned and unmanned low Earth orbit applications. The resultant nickel-hydrogen battery system should have a number of features that would lead to improved reliability, reduced costs as well as superior energy density and cycle lives as compared to battery systems constructed from the current state-of-the-art nickel-hydrogen individual pressure vessel cells.

Description: Partitioning of hydrogen chloride between the aerosol and gaseous phases in the first Space Shuttle exhaust cloud was experimentally investigated as the exhaust cloud was diluted with ambient air. Airborne measurements were obtained of gaseous hydrogen chloride (HCl), total HCl, relative humidity, and temperature to determine the conditions controlling HCl aerosol formation in the Shuttle exhaust cloud. Two segments of the cloud, each at a significantly different relative humidity, were monitored. Equilibrium predictions of HCl aerosol formation agreed with the measured HCl partitioning at the higher and lower relative humidity conditions, but do not agree at the aerosol formation threshold region. Measurements were taken in the Shuttle exhaust cloud from 8.6 min until 2 h and 8 min after launch. HCl concentrations ranged from 17.5 to 0.9 ppm and relative humidity from 86% to less than 10%.

Description: Single-stage and two-stage launch vehicles were evaluated for various levels of propulsion technology and payloads. The evaluation included tradeoffs between ascent flight performance and vehicle sizing that were driven by engine mass, specific impulse, and propellant requirements. Numerous mission, flight, and vehicle-related requirements and constraints were satisfied in the design process. The results showed that advanced technology had a large effect on reducing both single- and two-stage vehicle size. High-pressure hydrocarbon-fueled engines that were burned in parallel with two-position nozzle hydrogen-fueled engines reduced dry mass by 23% for the two-stage vehicle and 28% for the single-stage vehicle as compared to an all-hydrogen-fueled system. The dual-expander engine reduced single-stage vehicle dry mass by 41%. Using advanced technology, the single-stage vehicle became comparable in size and sensitivity to that of the two-stage vehicle for small payloads.

Description: Attention is directed to the acoustics research of the 1950s and 1960s for guidance in understanding and quantizing the turbulence amplification that can occur in regions of shock-wave boundary-layer interaction. Three primary turbulence amplifier-generator mechanisms are identified and shown, by linear analysis, to be responsible for turbulence amplification across a shock wave in excess of 100% of the incident turbulence intensity.

Description: A comparison is made between the storable-propellant engine and an equivalent dual-expander engine using oxygen, propane, and hydrogen propellants. The propane and hydrogen dual-expander engine is then compared with previous results with separate propane and hydrocarbon engines. It is shown that the dual-expander reduces vehicle dry mass from 94 to 89 Mg, or 5%. The hydrogen-only thrust of the dual-expander is 0.4 MN, while the separate hydrogen engines have a thrust of 2.3 MN. The low thrust level increases the difficulty of cooling the dual-expander engine, and additional hydrogen flow is added for transpirational cooling. This additional hydrogen flow leads to an increase in the hydrogen tank size. It is also shown that storable propellants increase the dry mass of single-stage earth-to-orbit vehicles by about 20% and the gross mass by about 54%. It is concluded that dual-expander engines should be studied with several thrust splits and thrust levels.

Description: An alternative propulsion subsystem for MSAT is presented which has a potential of reducing the satellite weight by more than 15%. The characteristics of pulsed plasma and ion engines are described and used to estimate of the mass of the propellant and thrusters for attitude control and stationkeeping functions for MSAT. Preliminary estimates indicate that the electric propulsion systems could also replace the large momentum wheels necessary to counteract the solar pressure; however, the fine pointing wheels would be retained. Estimates also show that either electric propulsion system can save approximately 18% to 20% of the initial 4,000 kg mass. The issues that require further experimentation are mentioned.

Description: Illustrations for a presentation on superplastic forming/diffusion bonding titanium design concepts are presented. Sandwich skin panels with hat section, semicircular corrugation, sine wave, and truss cores are shown. The fabrication of wing panels is illustrated, and applications to the design of advanced variable sweep bombers summarized.

Description: Illustrations for a presentation demonstrating superplastic forming/diffusion bonding titanium porous panels are presented. Fabrication phases, sandwich panels, load bearing qualities, microstructure, and panel surface after finishing are illustrated.

Description: The X-29A is a technology demonstrator. The FSW is just one of the technologies. Others include the following: discrete variable camber, relaxed static stability, triplex digital fly-by-wire (FBW) control system, variable-incidence/close-coupled canard, aeroelastically tailored composite wing, and thin supercritical airfoil. The growth potential for additional technologies is shown.

Description: The use of correlated data in airplane development is discussed. Areas of interest include initial airworthiness of an aircraft, low-speed configuration optimization, and high-speed configuration optimization. Data from wind tunnel tests are shown to be significant when applied to guarantee compliance, which includes fuel consumption, airspeeds, and takeoff and landing performance. The use of correlation in achieving FAA certification is also discussed.

Description: F-15 correlation data for longitudinal control and inlet-ramp effectiveness, and horizontal-tail setting for trim are presented. The Reynolds number effect on airfoil laminar bubble burst is included.

Description: The XB-70-1 was selected for a wind-tunnel/flight correlation program as representative of a large, flexible supersonic airplane similar to a supersonic transport. Tests were made to determine the effects of control deflections, wing tip deflection, and variations in inlet mass flow (additive drag).

Description: Priorties are identified for spacecraft propulsion system development and for the integration of the propulsion system with various subsystems. Near-term and long-term propulsion technology needs are identified.

Description: Orbital construction, demonstration, space construction system analysis, solar power satellite, and space operations are reviewed. Satellite services, holding and positioning aids, and space construction experiments are discussed.

Description: The workings of systems integration, its accomplishments, the influences of its character changes on the STS, propulsion out of the orbiter and LSS, and technological demands are discussed. The task of systems integration is to define, understand, and account for interactions between the major systems on a space mission. The safety and propulsion systems and their reliability are outlined.

Description: The functions and requirements of auxiliary propulsion systems are reviewed. None of the three major tasks (attitude control, stationkeeping, and shape control) can be performed by a collection of thrusters at a single central location. If a centralized system is defined as a collection of separated clusters, made up of the minimum number of propulsion units, then such a system can provide attitude control and stationkeeping for most vehicles. A distributed propulsion system is characterized by more numerous propulsion units in a regularly distributed arrangement. Various proposed large space systems are reviewed and it is concluded that centralized auxiliary propulsion is best suited to vehicles with a relatively rigid core. These vehicles may carry a number of flexible or movable appendages. A second group, consisting of one or more large flexible flat plates, may need distributed propulsion for shape control. There is a third group, consisting of vehicles built up from multiple shuttle launches, which may be forced into a distributed system because of the need to add additional propulsion units as the vehicles grow. The effects of distributed propulsion on a beam-like structure were examined. The deflection of the structure under both translational and rotational thrusts is shown as a function of the number of equally spaced thrusters. When two thrusters only are used it is shown that location is an important parameter. The possibility of using distributed propulsion to achieve minimum overall system weight is also examined. Finally, an examination of the active damping by distributed propulsion is described.

Description: Orbit transfer vehicles (OTV) are evaluated against mission model requirement. Mission suitability of OTVs using storable propellant, cryogenic propellant, and electric propulsion systems is outlined. Energy required, g-level, spacecraft deploy/return, operational constraints, mission duration, and packaging are considered.

Description: Possible uses of satellite technology up to the year 2000 are suggested and discussed. Included are electronic mail transmission, a personal communications capability, quick location of vehicles or shipments, monitoring of disputed territorial borders, upgraded scientific exploration of the universe, providing better maps of the Earth by remote sensing, space solar power stations and the safe transmission of the electrical energy to Earth, night lighting, and a small personal navigation capability. Support requirements are outlined.

Description: Considerations and recommendations for correlation are given. Basic tunnel calibration prior to research and development tests is suggested. Areas of concentration include: wing cruise drag and drag rise, wing separation and stall, afterbody and base drag, propulsion effects, vortex flows, cavity flows, and excrescences.

Description: A 2-percent-scale model was designed for testing in the NTF. This model has remotely controlled elevons, body flap, and rudder to minimize tunnel entries associated with configuration changes in the NTF. The Shuttle Orbiter has a very large aerodynamic data base obtained in ground facilities. Since the vehicle flight-test program has already begun, a large amount of flight data can be analyzed and correlated with the NTF results.

Description: It is found from the comparisons that large longitudinal aerodynamic differences exist between wind tunnel predictions and flight measurements. Cold gas plume simulation underpredicted Shuttle base pressure. It is concluded that observed flight prediction increments are probably caused by several factors such as input error, independent variable errors, plume effects, and Reynolds number effects.

Description: Correlation efforts and selected results for transonic drag are reviewed. A process to reduce the typical error sources to decrease the errors inherent in the transonic aircraft development process is summarized.

Description: Predicted and flight-test drag on the C-5A and the C-141 are correlated. Equivalent rigid flight-test profile drag and a rigid estimate based on wind tunnel data are also correlated. Correlations for the National Transonic Facility are included.

Description: Wind tunnel/flight correlation history from the P-51 to the F-8 supercritical wing is reviewed, showing that researchers continue to be faced with nearly identical discrepancies in predicted versus measured drag. The capabilities of the National Transonic Facility to allow assessment of the effects which have heretofore plagued researchers and aircraft designers are anticipated.

Description: The explicit-implicit predictor corrector method of MacCormack (1981) is applied to the analysis of flows past airfoils. By comparing results obtained with different methods and meshes, it is shown that the above method provides, after certain modifications, reasonably good predictions of inviscid and viscous flows about an airfoil. Good results are also obtained for the transonic regime if the free-stream conditions are correct and if a suitable mesh is used.

Description: Silicon solar cells manufactured for the terrestrial market are examined as a potential low cost option for low earth orbit (LEO) space flight use. The results of simulated space environmental testing of representative samples are reported and discussed. It is shown that although the terrestrial cells are compatible with most space use requirements significant deficiencies still exist. Cell modifications are discussed which would enhance the space applicability of the various cells examined. In most cases these are expected to be of minimal cost impact. Concern for contact/interconnector designs capable of surviving 30,000 thermal cycles (corresponding to five years in LEO) however, needs to be resolved for the large area terrestrial devices.

Description: An improved method, based on one strip approximation of the method of integral relations which was reported originally by Belov, Ginzburg and Shub (1973), is presented for the calculation of flow parameters in the impingement region of a supersonic, underexpanded jet striking a normal surface located within the first cell. The results are presented for two impingement conditions and found to be in good agreement with the experimental data.

Description: The Applications Technology Satellite-6 (ATS-6) geosynchronous satellite charged up to -2200 V in sunlight on day 178, 1974. This event, being the highest known spacecraft charging event in sunlight, is used to estimate a worst case geosynchronous plasma environment for predicting the spacecraft potential in eclipse. The advantage of using this sunlight spectrum as opposed to an eclipse case is that the ion and electron fluxes to the detectors are shifted only slightly due to the spacecraft potential. After correcting the available data for satellite potential and missing data above 81 KeV, it is found that the plasma can be characterized by a single Maxwellian approximation having an electron density of 1.22/cu cm, electron temperature of 16 KeV, hydrogen ion density of 0.24/cu cm, and hydrogen ion temperature of 29 KeV. In eclipse the spacecraft would have charged up to -28 kV, the highest estimated potential to date in the earth's plasma environment.

Description: A model is presented that explains the 'flat-spot' power-loss phenomenon observed in silicon solar cells operating under deep space (low temperature, low intensity) conditions. Evidence is presented suggesting that the effect is due to localized metallurgical interactions between the silicon substrate and the contact metallization. These reactions are shown to result in localized regions in which the PN junction is destroyed and replaced with a metal-semiconductor-like interface. The effects of thermal treatment, crystallographic orientation, junction depth, and metallization are presented along with a method of preventing the effect through the suppression of vacancy formation at the free surface of the contact metallization. Preliminary data indicating the effectiveness of a TiN diffusion barrier in preventing the effect are also given.

Description: Initial results of a NASA study of the lift-drag characteristics of a 12.84/7 deg biconic model intended for airbraking during atmospheric entry of probes to Mars, Venus, Saturn, and Titan are reported. Pressure distributions and shock shapes were measured in the Langley 20 in. Mach 6 tunnel with the spherically blunted bent-nose model set at angles from 0-25 deg. Pressure distributions and shock shapes where computed using the STEIN flowfield code, which features a MacCormack scheme to integrate the three-dimensional Euler equations, the Rankine-Hugoniot jump conditions to model shock waves as discontinuities, and requires a supersonic condition at every step. A comparison was made between measured and predicted values. The leeward shock angle was found to be predictable to within 3% for all angles of attack, while parabolized Navier-Stokes equations are regarded as offering more accurate results than the STEIN code for surface pressure distributions.

Description: The Lomax and Sluder method for adapting slender-wing theory to delta or rectangular wings by making chordwise and compressibility corrections is extended to cover wings of any arbitrary planform in subsonic and supersonic flows. The numerical accuracy of the present work is better than that of the Lomax-Sluder results. Comparison of the results of this work with those of the vortex-lattice method and Kernel function method for a family of Gothic and arrowhead wings shows good agreement. A universal curve is proposed for the evaluation of the lift coefficient of a low aspect ratio wing of an arbitrary planform in subsonic flow. The location of the center of pressure can also be estimated.

Description: (Previously announced in STAR as N81-30385)

Description: A procedure for the evaluation of wall interference corrections for three-dimensional aircraft configurations is presented. The Mach number and angle-of-attack corrections are obtained by numerically solving the Laplace equation in a parallelepiped with boundary conditions supplied mainly from experimental pressure measurements. A portion of these measurements and other wind-tunnel data required by the procedure may be replaced by theoretical estimates if not available from experiments. The accuracy of the correction results will then depend on the accuracy of these estimates. The correction procedure is applied to an isolated wing and to a wing-tail configuration in a solid-wall wind tunnel. It is found that neglecting twist and camber corrections for the wing effectively increases the tail angle-of-attack correction. Two different Mach number corrections can be calculated for the wing and tail. However, since only one Mach number correction is allowed for both the wing and the tail, and since the wing surface area is larger than the tail surface area, the final correction tends to be closer to the required wing correction. This is a source of error for the tail results.

Description: (Previously announced in STAR as N81-29092)

Description: Helicopter drop tests were made of models of the Pioneer Venus probe descent configurations to characterize their unsteady forces and angular dynamics in equilibrium descent. The axial and normal forces were found to be unsteady in magnitude by about 10 and 5% of the mean axial force, respectively. A cycle of the axial variation takes place in flight distances of from 15 to 40 diameters. The unsteadiness almost certainly is associated with the wake. Angular motions which do not converge to zero angle of attack even in very long duration descent are excited by the unsteady pitching moments. The nearly spherical large probe model was aerodynamically more unsteady than the round-nosed conical small probe model. Data returned from Venus by the Pioneer Venus probes show unsteady axial forces and angular motions similar to those seen in the drop tests.

Description: Cross-correlation techniques are used to measure the sound radiated by wing/flap airfoil configurations in the NASA-Ames 40 x 80 ft wind tunnel using a 6.7-m semispan model with three deployed flaps. The dominant source of flap noise is identified as the flap side edges, which exceeds that radiated by the midspan region by more than 10 dB. The turbulent surface eddies at the flap side edge have scales on the order of one-half the flap chord. The installation of flap actuator fairings at the flap side edge reduces the noise radiated from that location by 10 to 15 dB. The cross-correlation technique extracts airframe noise radiated by specific surface locations from the tunnel background noise, even when the noise is 25 dB higher than the measured airframe noise level.

Description: Computed solutions of the time-dependent, Reynolds-averaged Navier-Stokes equations for three-dimensional flows having thin shear layers are analyzed, using topological concepts. Specific examples include the transonic flow over a body of revolution with conical afterbody at moderate angles of incidence to the free stream. Experimental flow-visualization techniques are simulated graphically to visualize the computed flow. Scalar and vector fluid dynamic properties, such as pressure, shear stress, and vorticity on the body surface, are presented as topological maps, and their relationship to one another in terms of orientation and singular points is discussed. The extrapolation from these surface topologies toward the understanding of external flow-field behavior is discussed and demonstrated.

Description: The evaluation techniques, results and conclusions for the flight flutter testing conducted on a KC-135A airplane configured with and without winglets are discussed. Test results are presented for the critical symmetric and antisymmetric modes for a fuel distribution that consisted of 10,000 pounds in each wing main tank and empty reserve tanks. The results indicated that a lightly damped oscillation was experienced for a winglet configuration of a 0 deg cant and -4 deg incidence. The effects of cant and incidence angle variation on the critical modes are also discussed. Lightly damped oscillations were not encountered for any other winglet cant and incidence angles tested.

Description: One of the objectives of the KC-135 Winglet Flight Research and Demonstration Program was to obtain experimental flight test data to verify the theoretical and wind tunnel winglet aerodynamic performance prediction methods. Good agreement between analytic, wind tunnel and flight test performance was obtained when the known differences between the tests and analyses were accounted for. The flight test measured fuel mileage improvements for a 0.78 Mach number was 3.1 percent at 8 x 10(5) pounds W/delta and 5.5 percent at 1.05 x 10(6) pounds W/delta. Correcting the flight measured data for surface pressure differences between wind tunnel and flight resulted in a fuel mileage improvement of 4.4 percent at 8 x 10(5) pounds W/delta and 7.2 percent at 1.05 x 10(6) pounds W/delta. The performance improvement obtained was within the wind tunnel test data obtained from two different wind tunnel models. The buffet boundary data obtained for the baseline configuration was in good agreement with previous established data. Buffet data for the 15 deg cant/-4 deg incidence configuration showed a slight improvement, while the 15 deg cant/-2 deg incidence and 0 deg cant/-4 deg incidence data showed a slight deterioration.

Description: (Previously cited in issue 19, p. 3284, Accession no. A81-40931)

Description: The design of a high lift system for the NASA advanced LFC airfoil designed by Pfenninger is described. The high lift system consists of both leading and trailing edge flaps. A 3 meter semispan, 1 meter chord wing model using the above airfoil and high lift system is under construction and will be tested in the NASA Langley 4 by 7 meter tunnel. This model will have two separate full span leading edge flaps (0.10c and 0.12c) and one full span trailing edge flap (0.25c). The performance of this high lift system was predicted by the NASA two dimensional viscous multicomponent airfoil program. This program was also used to predict the characteristics of the LFC airfoils developed by the Douglas Aircraft Company and Lockheed-Georgia Aircraft Company.

Description: The design and construction of an advanced swept supercritical airfoil for commercial aircraft to be tested in a transonic wind tunnel is described. The swept LFC airfoil was designed for a given thickness ratio and lift coefficient, with emphasis placed on high critical Mach number with shock-free flow. It is compatible with satisfactory low speed and buffeting characteristics and minimizing the suction laminarization. Further emphasis was placed on achieving shock-free flow over a wide range of off-design conditions including trailing edge flap control. The requirements and design of the suction system and modifications to the Langley 8 foot transonic pressure tunnel is briefly described. Contouring of nonporous test section walls for free air simulation and flow quality improvements is included.

Description: The capability for calculating transonic flows for realistic configurations and conditions is discussed. Various phenomena which were modeled are shown to have the same order of magnitude on the influence of predicted results. It is concluded that CFD can make the following contributions to the task of correlating wind tunnel and flight test data: some effects of geometry differences and aeroelastic distortion can be predicted; tunnel wall effects can be assessed and corrected for; and the effects of model support systems and free stream nonuniformities can be modeled.

Description: A method for combining the cloud detector observation results from the Global Atmospheric Sampling Program (GASP) with Knollenberg probe observations of cloud particle concentration from other programs to derive estimates of the ambient concentration of particles larger than a given size was developed. The method was applied to estimate the probability of encountering particle concentrations which would degrade the performance of laminar flow control (LFC) aircraft. It is concluded that LF loss should occur only about one percent of the time in clear air and that flight within clouds should always result in a significant loss of LF, with 90 percent LF loss occurring about one percent of the time. Preliminary estimates of cloud encounter probability are presented for four airline routes, and conclusions are presented as to the best altitudes for cloud avoidance in extratropical and tropical latitudes.

Description: Some performance requirements and development needs for the design of large space structures are described. Areas of study include: (1) dynamic response of large space structures; (2) structural control and systems integration; (3) attitude control; and (4) large optics and flexibility. Reference is made to a large space telescope.

Description: The constraints placed on the design of large space structures by acceleration, attitude control, and stationkeeping forces are discussed. Stiffness requirements for the structures are derived. The use of active versus passive accuracy control methods is also addressed.

Description: The low Earth orbit (LEO) versus the geosynchronous Earth orbit (GEO) deployment systems were discussed. The following items are emphasized: (1) large area systems, such as deployment, altitude, orbit transfer, and on orbit operation; (2) propulsion oriented issues such as the orbit transfer vehicle and the auxiliary propulsion systems; (3) programmatic issues. The LEO versus GEO deployment is a significant driver on propulsion requirements for an orbit transfer vehicle. It is suggested that early systems will be deployed at LEO. The Shuttle remote manipulator system (RMS) will not be involved in early demonstration activities. It is concluded that an integrated large space system/propulsion approach is essential and that propulsion development requirements need to be established.

Description: The effects of low-thrust primary propulsion system characteristics on the mass, area, and orbit transfer characteristics of large space systems (LSS) were determined. Three general structural classes of LSS were considered, each with a broad range of diameters and nonstructural surface densities. While transferring the deployed structure from LEO and to GEO, an acceleration range of 0.02 to 0.1 g's was found to maximize deliverable payload based on structural mass impact. After propulsion system parametric analyses considering four propellant combinations produced values for available payload mass, length and volume, a thrust level range which maximizes deliverable LSS diameter was determined corresponding to a structure and propulsion vehicle. The engine start and/or shutdown thrust transients on the last orbit transfer (apogee) burn can impose transient loads which would be greater than the steady-state loads at the burnout acceleration. The effect of the engine thrust transients on the LSS was determined from the dynamic models upon which various engine ramps were imposed.

Description: Studies determined that shuttle optimized design, allowing the large space system and transfer vehicle in one shuttle flight, greatly reduces transportation costs and minimizes orbital operations. Careful attention to design resulted in efficient payload packaging. A minimum volume, high energy (liquid oxygen and liquid hydrogen) transfer vehicle is described that allows maximum volume for the payload in the orbiter cargo bay.

Description: Opportunities for improving the accuracy and reliability of wall corrections in conventional ventilated test sections are presented. The approach encompasses state-of-the-art technology in transonic computational methods combined with the measurement of tunnel-wall pressures. The objective is to arrive at correction procedures of known, verifiable accuracy that are practical within a production testing environment. It is concluded that: accurate and reliable correction procedures can be developed for cruise-type aerodynamic testing for any wall configuration; passive walls can be optimized for minimal interference for cruise-type aerodynamic testing (tailored slots, variable open area ratio, etc.); monitoring and assessment of noncorrectable interference (buoyancy and curvature in a transonic stream) can be an integral part of a correction procedure; and reasonably good correction procedures can probably be developd for complex flows involving extensive separation and other unpredictable phenomena.

Description: Correlation is defined in three different ways: comparisons of wind tunnel and/or theory with flight results; detailed studies of total vehicle drag from wind tunnel and flight tests; and attempts to understand the fundamental mechanisms of fluid flows associated with aircraft components in specific areas of the flight environment. The F-16E configuration is an outgrowth of studies conducted to produce a refined fighter wing design. Several iterations required to arrive at the combination of wing planform, camber, and twist which gives near optimum lift, drag, and high-angle-of-attack stability. Theoretical analyses were backed up by extensive experimental data to validate the design and are presented

Description: A conventional flight-test program, which slowly and cautiously approaches more severe flight conditions, was not possible with the Orbiter. On the first flight, the Orbiter entered the atmosphere at Mach 28 and decelerated through the Mach range. (The subsonic portion of flight was also flown by another orbiter vehicle during the Approach and Landing Test Program.) Certification for the first flight was achieved by an extensive wind-tunnel test and analysis program and by restricting the flight maneuvers severely. The initial flights of the orbiter were heavily instrumented for the purpose of obtaining accurate aerodynamic data. Even without maneuvers to excite the system, the first flight provided comparisons between flight and wind-tunnel-derived predicted data in the areas of aerodynamic performance, longitudinal trim, and reaction-control jet interaction. The aerodynamic performance comparisons are presented.

Description: Flow quality is discussed. Incremental comparisons of: (1) the angle of attack, (2) the axial force coefficient, and (3) the base cavity axial force coefficient against the normal force coefficient are presented. Relative blockage determination, relative buoyancy corrections, and boundary layer transition length are discussed. Blockage buoyancy caused by tunnel model wall dynamic interaction is discussed in terms of adaptive walls. The effect of 'transonic turbulence factor' is considered.

Description: Power systems integration in large flexible space structures is discussed with emphasis upon body control. A solar array is discussed as a typical example of spacecraft configuration problems. Information on how electric batteries dominate life-cycle costs is presented in chart form. Information is given on liquid metal droplet generators and collectors, hot spot analysis, power dissipation in solar arrays, solar array protection optimization, and electromagnetic compatibility for a power system platform.

Description: A rigid body analysis of a baseline Large Space System (LSS) which is to function as a radiometer is presented. The LSS is placed in circular orbit about the Earth at an altitude of 650 km, subjected to environmental and vehicle interaction forces and torques, without an active control system of any type on board. The environment forces and torques are gravity gradient, solar radiation, and aerodynamic. Normal operation is in nadir pointing along the Z-local vertical axis. Orbital velocity is assigned to the x-axis of the spacecraft. The analysis is then used to demonstrate the ability or lack of the gravity gradient torques to stabilize the LSS over one complete orbit.

Description: Requirements of future space systems, including large space systems, that operate beyond the space shuttle are discussed. Typical functions required of propulsion systems in this operational regime include payload placement, retrieval, observation, servicing, space debris control and support to large space systems. These functional requirements are discussed in conjunction with two classes of propulsion systems: (1) primary or orbit transfer vehicle (OTV) and (2) secondary or systems that generally operate within or relatively near an operational base orbit. Three propulsion system types are described in relation to these requirements: cryogenic OTV, teleoperator maneuvering system and a solar electric OTV.

Description: The dynamics of the interaction of space structures and their propellant systems are outlined. Optimization for a platform type of space structure is discussed. Static and transient loads, propellant accelerations, tolerances, attitude control, and distributed thrust are considered.

Description: Electric propulsion systems for transferring large payload masses to geosynchronous Earth orbits and providing accurate on-orbit stationkeeping are evaluated. Orbit boosting, inclination change, attitude control, stationkeeping, relocation, disposal, and power sharing on orbits using electric propulsion are compared with the use of chemical propulsion.

Description: A workshop panel defined missions that would require the potential use of very large, advanced space systems, and ranked them on the basis of need from both a civilian and military perspective. The panel also pointed out those that would need advanced propulsion technology. Surveillance, communications, and defense were given the highest priority, followed by command and control, orbital support, terrestrial support, and space science.

Description: Potential mission opportunities outlined and illustrated are missile defense; space defense; command, control, and communications; defense suppression; force support; space transportation; and orbital support.

Description: The overall technology model is outlined and the objectives and descriptions of the primary and secondary propulsion drives presented. The primary propulsion driver missions are the geostationary platform; the coherent optical system of modular imaging collectors (COSMIC); the 100 meter thinned aperature telescope; and the orbiting deep space relay station (ODSRS). The secondary propulsion driver missions are space platform alpha, the space station, and the automated planetary station.

Description: (Previously cited in issue 14, p. 2320, Accession no. A81-32905)

Description: A recently reported parabolized Navier-Stokes code has been employed to compute the supersonic flowfield about a spinning cone and spinning and nonspinning ogive cylinder and boattailed bodies of revolution at moderate incidence. The computations were performed for flow conditions where extensive measurements for wall pressure, boundary-layer velocity profiles, and Magnus force had been obtained. Comparisons between the computational results and experiment indicate excellent agreement for angles of attack up to 6 deg. At angles greater than 6 deg discrepancies are noted which are tentatively attributed to turbulence modeling errors. The comparisons for Magnus effects show that the code accurately predicts the effects of body shape for the selected models.

Description: (Previously cited in issue 17, p. 3111, Accession no. A80-41562)

Description: Conventional methods for the calculation of wall interference corrections are based on boundary conditions which require a knowledge of ventilated wall porosity parameters, and which are unsuitable for deformed walls. The method described uses a simple exponential decay of pressure distribution beyond the most upstream and downstream limits in order to evaluate Mach number and incidence corrections given by the method proposed by Papelier et al. (1978). It is found that, while the upstream contribution to incidence correction is significant, the upstream and downstream contributions to Mach number correction are negligible.

Description: (Previously cited in issue 07, p. 963, Accession no. A82-19212)

Description: An exact analytical solution for the velocity field, both interior and exterior, induced by an infinite right-handed helical vortex filament is derived. Due to the way the variables combine in this solution, the paper also shows that it is possible to derive a stream function for this nonaxisymmetric flow. Sample calculations of these expressions are included.

Description: Airborne measurements of two types of cloud nuclei, cloud condensation nuclei (CCN) and ice nuclei (IN), were conducted in the stabilized ground clouds which resulted from the launches of a liquid-fueled ATLAS/Centaur rocket and a solid-fueled TITAN III rocket. Results show that the concentrations of CCN in both types of clouds were greater than ambient values for the 2 hours duration of the measurements. The initial production of CCN active at 0.5% supersaturation in the ATLAS and TITAN clouds was found to be equivalent to a 20 and 700 sec emission, respectively, by the city of Denver, Colorado. After the initial production, the clouds continued to generate CCN at a rate of about 1/cu cm sec. However, concentrations of IN in the ATLAS cloud were greater than ambient values for only a short period after launch, and it appears that the nuclei resulted from entrained launch pad and ground debris. The concentrations of IN in the TITAN cloud were found to be at or below ambient levels until about 2 hours after launch when they increased substantially above ambient values.

Description: Design concepts, study results, and research directions toward development of CW laser heating of remotely flying spacecraft fuels to provide high impulse thrust are presented. The incident laser radiation would be absorbed by hydrogen through a medium of a laser-supported plasma. The laser energy could be furnished from an orbiting solar-powered laser platform and used to drive the engines of an orbital transfer vehicle (OTV) at costs less than with a chemical propulsion system. The OTV propulsion chamber would be reduced in size comparable to the volume addition of the incident laser energy absorber. The temperatures in the hydrogen-fueled system could reach 5000-15,000 K, and studies have been done to examine the feasibility of ion-electron recombination. Kinetic performance, temperature field, and power necessary to sustain a laser thrust augmented system modeling results are discussed, along with near-term 30 kW CO2 laser system tests.

Description: Analytical studies and laboratory experiments have been performed to evaluate the vibration response of the Space Shuttle Thermal Protection System (TPS) tiles due to the intense rocket generated acoustic noise during lift-off. The TPS tiles are mounted over the exterior of the Space Shuttle Orbiter structure through Strain Isolation Pads (SIP) which protect the tiles from thermal induced shear loads at their interface. The analytical predictions indicate that the response of a typical tile is governed by the structural vibration inputs through the SIP under the tile at frequencies below 250 Hz, and by the direct acoustic excitation over the exterior surface of the tile at frequencies above 250 Hz. An evaluation of the laboratory test data for this same tile, in which conditioned (partial) coherent output spectral analysis procedures were used, leads to exactly the same conclusion. The results demonstrate the power of conditioned spectral analysis procedures in identifying vibration response mechanisms when two or more of the inputs are highly correlated.

Description: Rapid double-pulse holography was employed to obtain detailed, two-dimensional images of the shock forming during simulated flutter in a transonic flowfield. The experiment comprised a linear cascade of airfoils externally oscillated in torsion and viewed tangentially at the shock surface. Three biconvex airfoils were subjected to harmonic pitching motion about the midchord axis at a frequency of 0.53 while immersed in a Mach 0.81 flow. Failure to produce observable shocks led to use of choked flow with a Mach number near one, of which 50 holograms were taken. The images revealed a narrow shock surface with a spanwise variation in the shock properties. The method is concluded to be useful for examining transonic flowfield shocks in the presence of airfoil flutter.

Description: Progress towards successful modelling of unsteady flows past two-dimensional oscillating airfoils is examined. Linearized, thin airfoil theory is reviewed with special regard to the vortex shedding which occurs when either the body or the flow fluctuates. A sinusoidally oscillating flat-plate airfoil is considered in terms of noncirculatory components, including boundary conditions and bound and wake vortices, respectively. Applications of linear theory to vertical airfoil oscillations and to oscillating control surfaces are described, and oscillating airfoils in subsonic and supersonic flows are investigated. Perturbations in linear solutions are explored for the occurrence of second-order effects, and trailing edge and wake effects. The effects of unsteady transonic flows are broken into nearly inviscid and strongly viscous cases, and analyses of dynamic stall and stall flutter are presented.

Description: Noise emission from very small chord and very large chord airfoils was measured with eleven 0.63 cm microphones placed along a horizontal semicircle (4.57 m radius) that was centered at the leading edge of the test airfoil. The noise signals were analyzed by an automated spectrum analyzer which yielded 1/3-octave band sound pressure level spectra for each microphone, and the data were corrected to remove the effects of atmospheric attenuation and jet noise. It is found that the effect of thickness is large and must be accounted for in any fundamental airfoil noise theory that attempts to describe the noise emitted from real airfoils. Incident mean velocity gradients and compressibility must also be taken into account. The effect of thickness increases with frequency, with thick airfoils being quieter than thin ones.

Description: A simple practical method for designing antenna-feed positioning control systems for large deployable spaceborne antenna systems with flexible booms is proposed. The approach is based on the mechanical decoupling of the antenna-feed from the boom so that the positioning control system can be designed without taking boom dynamics into consideration, thus avoiding a complex infinite dimensional control problem. The basic idea is illustrated by a simple angular positional control system attached to a flexible boom restricted to torsional motion only. The application of this approach to more complex situations is discussed briefly.