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Concentric glass-tube-envelope device surrounds flat-plate absorber having spectrally selective coating. Transparent envelope has antireflection coating. Heat-transfer medium is gas that circulates along hairpin path.
Four combustor concepts, designed for the energy efficient engine, utilize variable geometry or other flow modulation techniques to control the equivalence ratio of the initial burning zone. Lean conditions are maintained at high power to control oxides of nitrogen while near stoichometric conditions are maintained at low power for low CO and THC emissions. Each concept was analyzed and ranked for its potential in meeting the goals of the program. Although the primary goal of the program is a low level of nitric oxide emissions at stratospheric cruise conditions, both the ground level EPA emission standards and combustor performance and operational requirements typical of advanced subsonic aircraft engines are retained as goals as well. Based on the analytical projections made, two of the concepts offer the potential of achieving the emission goals; however, the projected operational characteristics and reliability of any concept to perform satisfactorily over an entire aircraft flight envelope would require extensive experimental substantiation before engine adaptation can be considered.
An arc-heated wind-tunnel system described in the present paper will simulate aerodynamic heating in large-scale tests of the thermal protection system of the Shuttle Orbiter Vehicle during entry. The system provides for large-scale subsystem tests in high-enthalpy streams with boundary layer flows at high Reynolds numbers and for large test-body sizes in stagnation flows. The discussion covers the design concept of the arc-jet systems, the extensive hardware developments of the arc heater to provide reliable operation, and verification and performance measurements of the system's operating envelopes.
The 30 cm J Series Thruster has been base-lined for the Solar Electric Propulsion Vehicle. This paper discusses the evolution of the thruster and details its performance, lifetime, and operational characteristics. An operational envelope covering nominal input powers of 2650 to 700 W is defined. Total thruster efficiencies over this envelope range from 0.711 to 0.472 with a corresponding impulse and thrust ranging from 2980 to 1850 sec and 129.1 to 36.4 mN, respectively. Completed life test segments and the ability of the thruster to meet the design goal of 15,000 hr of life as well as the life limitations are discussed. A description and rationale of the present set of thruster operational control algorithms is also presented.
Pressure data at 90 percent blade radius were obtained for a helicopter main rotor with 10-64C blade sections during flight. Concurrent measurements ere made of vehicle flight state, performance and some rotor loads. The test envelope included hover, level flight from about 65 to 162 knots, climb and descent, and collective fixed maneuvers. Good agreement is shown between some sets of airfoil pressure distributions obtained in flight and those from two-dimensional wind-tunnel tests or theoretical calculations.
Pressure data at 90 percent blade radius for a helicopter main rotor with RC-SC2 blade sections was obtained. Concurrent measurements were made of vehicle flight state, performance and some rotor loads. The test envelope included hover, level flight from about 65 to 144 knots, climb and descent, and collective fixed maneuvers. Airfoil pressure distributions obtained in flight agree with those theoretical calculations for two dimensional, steady flow.
Development of the Body Flap Actuation Subsystem for Space Shuttles included alterations from the original design to mechanical stops, planet gears, control valves, and solenoid valves. The mechanical stops were redesigned to absorb stall load and rotating inertia of the hydraulic motors instead of only stall load. The institution of a quill shaft (torsion spring) was a successful solution. The planet gears in the geared rotary actuators developed cracks during testing. This failure was alleviated via modification to the gears. A motor pressurization - brake release timing technique was developed thru analysis and testing. This resulted in a control valve configuration which would not permit freewheeling of the body flap surface. Finally, several solenoid valve configurations were tested to obtain the desired performance. Conceptual redesigns and modifications were weighted against each other to optimize a solution. Tradeoffs were usually made between life, performance, failure tolerance, and reliability versus weight, envelope, and maintainability.
An eclectic approach was taken to a study of the integration of digital flight and propulsion controls for helicopters. The basis of the evaluation was the current Gen Hel simulation of the UH-60A Black Hawk helicopter with a model of the GE T700 engine. A list of flight maneuver segments to be used in evaluating the effectiveness of such an integrated control system was composed, based on past experience and an extensive survey of the U.S. Army Air-to-Air Combat Test data. A number of possible features of an integrated system were examined and screened. Those that survived the screening were combined into a design that replaced the T700 fuel control and part of the control system in the UH-60A Gen Hel simulation. This design included portions of an existing pragmatic adaptive fuel control designed by the Chandler-Evans Company and an linear quadratic regulator (LQR) based N(p) governor designed by the GE company, combined with changes in the basic Sikorsky Aircraft designed control system. The integrated system exhibited improved total performance in many areas of the flight envelope.
A generic, state-of-the-art, high-performance aircraft model, including detailed, full-envelope, nonlinear aerodynamics, and full-envelope thrust and first-order engine response data is described. While this model was primarily developed Controls Design Challenge, the availability of such a model provides a common focus for research in aeronautical control theory and methodology. An implementation of this model using the FORTRAN computer language, associated routines furnished with the aircraft model, and techniques for interfacing these routines to external procedures is also described. Figures showing vehicle geometry, surfaces, and sign conventions are included.
The X-29A forward-swept wing flight research aircraft flight envelope was expanded to 66 deg angle-of-attack during 1990. Following this flight envelope expansion, a military utility evaluation was performed to investigate the tactical utility of the X-29 configurations at high-angle -of-attack, slow-speed flight conditions. An overall management view and perspective of the expansion process, the technical problems encountered, and the results obtained when compared to the predictions are presented.
A flight control concept that can identify aircraft stability properties and continually optimize the aircraft flying qualities has been developed by McDonnell Aircraft Company under a contract with the NASA-Dryden Flight Research Facility. This flight concept, termed the Intelligent Flight Control System, utilizes Neural Network technology to identify the host aircraft stability and control properties during flight, and use this information to design on-line the control system feedback gains to provide continuous optimum flight response. This self-repairing capability can provide high performance flight maneuvering response throughout large flight envelopes, such as needed for the National Aerospace Plane. Moreover, achieving this response early in the vehicle's development schedule will save cost.
The objective of the Innovative High Temperature Aircraft Engine Fuel Nozzle Program was to design and evaluate a nozzle capable of operating at a combustor inlet air temperature of 1600 F (1144 K) and a fuel temperature of 350 F (450 K). The nozzle was designed to meet the same performance requirements and fit within the size envelope of a current production dual orifice fuel nozzle. The design approach was to use improved thermal protection and fuel passage geometry in combination with fuel passage surface treatment to minimize coking at these extreme fuel and air temperatures. Heat transfer models of several fuel injector concepts were used to optimize the thermal protection, while sample tube coking tests were run to evaluate the effect of surface finish, coatings and tube material on the coking rate. Based on heat transfer analysis, additional air gaps, reduced fuel passage flow area and use of ceramic tip components reduced local fuel wetted wall temperatures by more than 200 F (100 K) when compared to a current production fuel nozzle. Sample tube coking test results showed the importance of surface finish on the fuel coking rate. Therefore, a one micro-inch (0.025 micron) roughness was specified for all fuel passage surfaces. A novel flow divider valve in the tip was also employed to reduce weight, allow room for additional thermal protection, and provide back pressure to reduce the risk of fuel vaporization. The fuel nozzle was fabricated and evaluated in a series of high temperature coking tests. Initial results of these tests indicate that thermal protection and surface treatments were partially successful in preventing nozzle fouling, but additional refinement of the internal flowpath is needed to prevent buildup of coke particles that partially blocked the spin slots which meter the nozzle fuel flow.
Advanced propulsion research and technology concepts require launch and space flight technologies, which can drastically reduce mission costs. Laser propulsion is a concept in which energy of a thrust producing reaction mass is supplied via beamed energy from an off-board power source. A variety of laser beamed energy concepts were theoretically and experimentally investigated since the early 1970's. During the 1980's the Strategic Defense Initiative (SDIO) research lead to the invention of the Laser Lightcraft concept. Based upon the Laser Lightcraft concept, the propulsion research center, TD 40, within the Space Transportation Directorate of NASA MSFC has set out to develop technologies required for launching small payloads into Earth Orbit for a cost of $1.0M or $1000/lb to $100/lb. The near term objectives are to demonstrate technologies and capabilities essential for a future earth to orbit launch capability. Laser propulsion offers the advantages of both high thrust and good specific impulse, I(sp) in excess of 1000 s(exp -1). Other advantages are the simplicity and reliability of the engine because of few moving parts; simplistic propellant feed system, and low specific fuel consumption as a result of high specific impulse. Major limitations of this approach are the laser power available, absorption and distortion of the pulsed laser beam through the atmosphere, and coupling laser power into thrust throughout the flight envelope. This summer's effort focused on performance of the laser engine, which included 1) optical ray tracing and 2), a time dependent calculation of the optically induced blast wave.
A NASA-wide team, funded under the NASA Electronic Parts and Packaging Program (NEPP), was formed to collaborate and to establish reliability of various electronic parts/packaging and assemblies for operation under extreme cold temperatures. One of the primary objectives of the NEPP is to expedite the infusion of cutting edge technologies into the present and future NASA missions. Commercial-off-the-shelf (COTS) emerging electronic parts/packages due to their lower weight, increased functionality, and lower cost are excellent candidates for space missions if they are characterized to show that they will meet the stringent reliability and quality requirements. Characterizations, especially for the extreme cold temperatures, are required since very limited data are available by manufacturers or users. For severe military environments, the temperature conditions to -65 C are the lowest temperature for which these parts/packages and assemblies are qualified. New data beyond this relatively benign cold temperature are required for numerous NASA missions. Several parts/packages, based on the project recommendation for their immediate and future needs, were selected for detailed characterization to cold temperature regimes down to liquid nitrogen (-196 C), covering both Mars cold temperature (-125 C) and asteroid (- 180 C) lander environmental requirements. Numerous parts/packages and assemblies were characterized during extreme temperature environmental tests. Several electrical parameters were characterized at discrete temperatures to -185 C to determine if they remain within their specification ranges. Both packages and circuit boards were subjected to nondestructive testing including optical, X-ray, and acoustic microscopy to document their integrity prior to environment exposure. Package/board assemblies were also subjected to X-ray to characterize solder joint integrity including void levels. Both parts and assemblies were subjected to thermal cycling with a large temperature swing enveloping numerous NASA missions. Details of the performed tests and the results obtained are presented.
Fan designs are often constrained by envelope, rotational speed, weight, and power. Aerodynamic performance and motor electrical performance are heavily influenced by rotational speed. The fan used in this work is at a practical limit for rotational speed due to motor performance characteristics, and there is no more space available in the packaging for a larger fan. The pressure rise requirements keep growing. The way to ordinarily accommodate a higher DP is to spin faster or grow the fan rotor diameter. The invention is to put two radially oriented stages on a single disk. Flow enters the first stage from the center; energy is imparted to the flow in the first stage blades, the flow is redirected some amount opposite to the direction of rotation in the fixed stators, and more energy is imparted to the flow in the second- stage blades. Without increasing either rotational speed or disk diameter, it is believed that as much as 50 percent more DP can be achieved with this design than with an ordinary, single-stage centrifugal design. This invention is useful primarily for fans having relatively low flow rates with relatively high pressure rise requirements.
Overall, the MSL TDS Field Test campaign was very successful. TDS was shown to perform extremely well over the required operational envelope. Early BB TDS field tests uncovered a number of issues, but none that invalidated the TDS design or implementation. EM TDS tests uncovered minor things of interest, but nothing of concern. Value of testing hardware in the field was demonstrated and significantly contributed to the overall TDS V&V effort. Over the 5-plus year field test campaign, numerous lessons were learned that will inform future field test efforts.
As part of the Stratospheric Observatory for Infrared Astronomy (SOFIA) program, a 747SP airplane was modified to carry a 2.5 meter telescope in the aft section of the fuselage. The resulting airborne observatory allows for observations above 99 percent of the water vapor in the atmosphere. The open cavity created by the modifications had the potential to significantly affect the airplane in the areas of aerodynamics and acoustics. Several series of flight tests were conducted to clear the airplanes operating envelope for astronomical observations, planned to be performed between the altitudes of 39,000 feet and 45,000 feet. The flight tests were successfully completed. Cavity acoustics were below design limits, and the overall acoustic characteristics of the cavity were better than expected. The modification did have some effects on the stability and control of the airplane, but these effects were not significant. Airplane air data systems were not affected by the modifications. This paper describes the methods used to examine the aerodynamics and acoustic data from the flight tests and provides a discussion of the flight test results in the areas of cavity acoustics, stability and control, and air data.
Experimental results are presented for a series of experiments that addressed the effect of small pinhole defects on the potential lifetime of a Venus superpressure balloon. The experiments were performed on samples of a candidate balloon envelope material through which a single small hole of 80 to 300 microns in diameter was deliberately made in each one by puncturing with a metal pin. The material was mounted horizontally in a test apparatus and then a 2-3 mm thick layer of sulfuric acid was placed on top to mimic balloon wetting at Venus. Acid penetration and damage manifested itself as a darkening of the aluminum metal and adhesive layers around the hole in the balloon material. There were no test conditions under which the acid simply fell through the pinhole due to gravity because the surface tension forces always compensated at this size. Very little acid-damaged material was observed for the smallest 80 micron pinholes while gas flowed through the hole due to balloon-like pressurization: the black spot size was approximately 0.2 mm in diameter after 6 days with 86% sulfuric acid. The damage area grew more quickly in the absence of gas flowing out of an 80 micron hole, namely at a rate of 2 mm/day. It was concluded that the flow of escaping gas out of the hole provides a substantial reduction of the rate of acid penetration and damage. Larger diameter pinholes of approximately 300 micron diameter showed larger growth rates of 0.7 mm/day with gas flow and 1.7 mm/day without. The pinhole size did not change over the duration of these experiments because the material has an outer layer of fluoropolymer film that remained intact during the process and thereby held the hole size constant. None of the damage rates measured in these experiments pose a threat to the lifetime of the balloon over the projected course of a 30 day mission because the affected area is too small to cause a structural failure either through direct damage or increased solar heating and attendant balloon pressurization leading to burst.