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Data on the performance, stability, and thermal characteristics of an OME operating with an alternate injector configuration and with alternate propellants was obtained. The design, manufacturing, and operating characteristics of an electroformed, regeneratively cooled thrust chamber were also derived. Subscale and full scale tests provide data relating to off-design and transient operation.
A series of discrete maneuver tasks were used to evaluate the effects of steering gain and directional mode dynamic parameters on driver/vehicle responses. The importance and ranking of these parameters were evaluated through changes in subjective driver ratings and performance measures obtained from transient maneuvers such as a double lane change, an emergency lane change, and an unexpected obstacle. The unexpected obstacle maneuver proved more sensitive to individual driver differences than to vehicle differences. Results were based on full scale tests with an experienced test driver evaluating many different dynamic configurations plus seventeen ordinary drivers evaluating six key configurations.
This paper presents the design development of a truss-type structure which utilizes graphite/epoxy as the primary material system. A review is given of an on-going NASA sponsored program to develop a structural assembly which is required to have a very high degree of dimensional stability. Methods of material characterization and selection are given along with property data such as strength, stiffness, outgassing and thermal expansion. Finite element (NASTRAN) analysis will be discussed and deflection plots of the lowest mode frequencies will be presented. Proposed full scale tests: static, dynamic, and thermal/vacuum will be described.
The feasibility of damping more than one mode of rocket engine combustion instability by means of differently tuned acoustic cavities sharing a common entrance was shown. Analytical procedures and acoustic modeling techniques for predicting the stability behavior of acoustic cavity designs in hot firings were developed. Full scale testing of various common entrance, dual cavity configurations, and subscale testing for the purpose of obtaining motion pictures of the cavity entrance region, to aid in determining the mechanism of cavity damping were the two major aspects of the program.
The article discusses the development and capabilities of the National Transonic Facility which is planned for operation in 1981. The fan-drive, cryogenic-pressurized, closed-return facility will have operating parameters of: 0.1-1.2 Mach, 1-9 bars pressure, 78-340 K, 150 dB sound pressure, and plus or minus 0.001 rms turbulence intensity. These operating conditions have been selected on the basis of several current and future aircraft and space transportation systems. The facility will provide full-scale testing conditions for calculating subsonic drag, airloads, and stability and control information. Data for pre-test conditions, on-line information, and post-test analysis will be computer-processed.
A study was conducted to define the technology and data needed to support the introduction of advanced composite materials in the wing structure of future production aircraft. The study accomplished the following: (1) definition of acceptance factors, (2) identification of technology issues, (3) evaluation of six candidate wing structures, (4) evaluation of five program options, (5) definition of a composite wing technology development plan, (6) identification of full-scale tests, (7) estimation of program costs for the total development plan, (8) forecast of future utilization of composites in commercial transport aircraft and (9) identification of critical technologies for timely program planning.
A brief review of LERC (Limiting Energy Release Criteria) is presented. Application of LERC to JSC/Boeing ignition source full-scale tests and to JSC/Dacfir math-model validation tests is outlined. It is concluded that: (1) a complete LERC application to the JSC/Boeing tests verifies the fuel load criterion as the limiting constraint; (2) the variability of magnitude and form of the results of repeated tests with and without small variations in parameters emphasizes the significance of the local flow, species concentration, and heat transfer distribution; and (3) weight loss measurements of recent JSC tests show consistent results with prior methods, fuel load constrained.
Acoustic data were obtained during a full-scale test of the XV-15 Tilt Rotor Research Aircraft in the Ames 40- by 80-Foot Wind Tunnel. The XV-15 has two 25-ft-diameter, three-bladed rotors at the tips of a 32-ft span wing. The rotors are used as lifting rotors, as propellers, and in various intermediate stages. Acoustic waveforms, dBA, and spectra as functions of different rotor conditions are presented and discussed. The noise level was found to be sensitive to change in rotor-disk angle of attack in the helicopter configuration. Much higher noise levels and harmonic contents were found in the helicopter mode than in the airplane mode. The measurements are compared with existing theoretical predictions which include wind-tunnel wall reflections.
An advanced composite frame which is flight-weight and integrates the functions of several structures was developed for the over the wing (OTW) engine and for the under the wing (UTW) engine. The composite material system selected as the basic material for the frame is Type AS graphite fiber in a Hercules 3501 epoxy resin matrix. The frame was analyzed using a finite element digital computer program. This program was used in an iterative fashion to arrive at practical thicknesses and ply orientations to achieve a final design that met all strength and stiffness requirements for critical conditions. Using this information, the detail design of each of the individual parts of the frame was completed and released. On the basis of these designs, the required tooling was designed to fabricate the various component parts of the frame. To verify the structural integrity of the critical joint areas, a full-scale test was conducted on the frame before engine testing. The testing of the frame established critical spring constants and subjected the frame to three critical load cases. The successful static load test was followed by 153 and 58 hours respectively of successful running on the UTW and OTW engines.
The development of an analytical model for the prediction of sound levels in the payload bay of the space shuttle orbiter vehicle is outlined. Formulation of the analytical model and its validation by means of model scale and full scale tests are included. It is shown that the approach used in the development effort has resulted in a prediction procedure which can be expected to give reliable estimates of payload bay sound levels, even when a payload is present. Furthermore, the analytical model has the capability of being readily modified to include other excitations such as turbulent boundary layers and propeller near-field pressures, and to other aerospace vehicles.
The paper considers the bidirectional reflectance distribution function (BRDF) of black coatings used on stray light suppression systems for the Space Telescope (ST). The ST stray light suppression requirement is to reduce earth, moon, and sun light in the focal plane to a level equivalent to one 23 Mv star per square arcsecond, an attenuation of 14 orders of magnitude. It is impractical to verify the performance of a proposed baffle system design by full scale tests because of the large size of the ST, so that a computer analysis is used to select the design. Accurate computer analysis requires a knowledge of the diffuse scatter at all angles from the surface of the coatings, for all angles of incident light. During the early phases of the ST program a BRDF scanner was built at the Marshall Space Flight Center to study the scatter from black materials; the measurement system is described and the results of measurements on samples proposed for use on the ST are presented.
Acoustic data were obtained during a full-scale test of the XH-59A Advancing Blade Concept Technology Demonstrator in the 40- by 80-Foot Wind Tunnel. The XH-59A is a research helicopter with two coaxial rotors and hingeless blades. Performance, vibration, and noise at various forward speeds, rotor lift coefficients and rotor shaft angles of attack were investigated. The noise data were acquired over an isolated rotor lift coefficient range of 0.024 to 0.162, an advance ratio range of 0.23 to 0.45 corresponding to tunnel wind speeds of 89 to 160 knots, and angles of attack from 0 deg to 10 deg. Acoustic data are presented for seven microphone locations for all run conditions where the model noise is above the background noise. Model test configuration and performance information are also listed. Acoustic waveforms, dBA, and 1/3-octave spectra as functions of operating condition for selected data points and microphones are presented. In general, the noise level is shown to increase with rotor lift coefficient except under certain operating conditions where significant impulsive blade/vortex interactions increase noise levels.
Acoustic data were obtained during a full-scale test of the XH-59A Advancing Blade Concept (ABC) Technology Demonstrator in the NASA Ames 40- by 80-Foot Wind Tunnel. The XH-59A is a research helicopter with two coaxial rotors and hingeless blades. Performance, vibration, noise at various forward speeds, rotor lift coefficients, and rotor shaft angles of attack were investigated. In general, the noise level is shown to increase with rotor lift coefficient except under certain operating conditions where it is increased by significant impulsive blade/vortex interactions. The impulsivity appears to depend upon how the lift is distributed between the two rotors. The noise levels measured are shown to be slightly higher than on a modern conventional rotor tested in the same facility.
Future spacecraft are being conceived that are highly flexible and of extreme size. The two features of flexibility and size pose new problems in control system design. Since large scale structures are not testable in ground based facilities, the decision on component placement must be made prior to full-scale tests on the spacecraft. Control law research is directed at solving problems of inadequate modelling knowledge prior to operation required to achieve peak performance. Another crucial problem addressed is accommodating failures in systems with smart components that are physically distributed on highly flexible structures. Parameter adaptive control is a method of promise that provides on-orbit tuning of the control system to improve performance by upgrading the mathematical model of the spacecraft during operation. Two specific questions are answered in this work. They are: What limits does on-line parameter identification with realistic sensors and actuators place on the ultimate achievable performance of a system in the highly flexible environment? Also, how well must the mathematical model used in on-board analytic redundancy be known and what are the reasonable expectations for advanced redundancy management schemes in the highly flexible and distributed component environment?
The vortex trajectory and vortex wake generated by helicopter rotors were visualized using a wide-field shadowgraph technique. Use of a retro-reflective Scotchlite screen made it possible to investigate the flow field generated by full-scale tail rotors. Tip vortex trajectories were visible in shadowgraphs for a range for tip Mach number of 0.38 - 0.60. The effect of the angle of attack was substantial. At an angle of attack greater than 8 degrees, the visibility of the vortex core was significant even at relatively low tip Mach numbers. The theoretical analysis of the visibility was carried out for a rotating blade. This analysis demonstrated that the visibility decreases with increasing dimensionless core radius (ro/c) and increases with increasing tip Mach number. Based on this investigation, it is concluded that the wide-field shadowgraph flow visualization technique should be feasible to study the flow field generated by a large main rotor in a wind tunnel and in an outdoor full-scale test stand. Of note is that the shadowgraph technique could easily be used with other on-going tests such as aerodynamic performance study, noise measurements, velocity measurements with LDV or hot-wire anemometer, local pressure measurement on the rotor surface, etc.
The design of a helicopter is based on an understanding of many parameters and their interactions. For example, in the design stage of a helicopter, the weight, engine, and rotor speed must be considered along with the rotor geometry when considering helicopter operations. However, the relationship between the noise radiated from the helicopter and these parameters is not well understood, with only limited model and full-scale test data to study. In general, these data have shown that reduced rotor speeds result in reduced far-field noise levels. This paper reviews the status of a recent helicopter noise research project designed to provide experimental flight data to be used to better understand helicopter rotor-speed effects on far-field acoustic levels. Preliminary results are presented relative to tests conducted with a McDonnell Douglas model 500E helicopter operating with the rotor speed as the control variable over the range of 103% of the main-rotor speed (NR) to 75% NR, and with the forward speed maintained at a constant value of 80 knots.
A study was conducted to define the technology and data needed to support the introduction of advanced composites in the future production of fuselage structure in large transport aircraft. Fuselage structures of six candidate airplanes were evaluated for the baseline component. The MD-100 was selected on the basis of its representation of 1990s fuselage structure, an available data base, its impact on the schedule and cost of the development program, and its availability and suitability for flight service evaluation. Acceptance criteria were defined, technology issues were identified, and a composite fuselage technology development plan, including full-scale tests, was identified. The plan was based on composite materials to be available in the mid to late 1980s. Program resources required to develop composite fuselage technology are estimated at a rough order of magnitude to be 877 man-years exclusive of the bird strike and impact dynamic test components. A conceptual composite fuselage was designed, retaining the basic MD-100 structural arrangement for doors, windows, wing, wheel wells, cockpit enclosure, major bulkheads, etc., resulting in a 32 percent weight savings.
The flow entraining capabilities of the Circulation Control Wing high lift system were employed to provide an even stronger STOL potential when synergistically combined with upper surface mounted engines. The resulting configurations generate very high supercirculation lift in addition to a vertical component of the pneumatically deflected engine thrust. A series of small scale wind tunnel tests and full scale static thrust deflection tests are discussed which provide a sufficient data base performance. These tests results show thrust deflections of greater than 90 deg produced pneumatically by nonmoving aerodynamic surfaces, and the ability to maintain constant high lift while varying the propulsive force from high thrust recovery required for short takeoff to high drag generation required for short low speed landings.