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The five basic elements affecting runway traction for jet transport aircraft operation are identified and described in terms of pilot, aircraft system, atmospheric, tire, and pavement performance factors or parameters. Where possible, research results are summarized, and means for restoring or improving runway traction for these different conditions are discussed.
Assembly of cylinders, links, actuators, and gears permits landing-gear unit to be retracted into shadow of main engine intake ducts of supersonic transport aircraft. This is accomplished without adding to frontal area of aircraft or appreciably increasing total aircraft drag.
Factors influencing crew and passenger exposure to solar and galactic cosmic rays that is expected to occur during flights of supersonic transport aircraft are discussed, and some possibilities are considered for decreasing such exposure.
Linear optimal regulator theory is applied to a nonlinear simulation of a transport aircraft performing a helical landing approach. A closed form expression for the quasi-steady nominal flight path is presented along with the method for determining the corresponding constant nominal control inputs. The Jacobian matrices and the weighting matrices in the cost functional are time varying. A method of solving for the optimal feedback gains is reviewed. The control system is tested on several alternative landing approaches using both three and six degree flight path angles. On each landing approach, the aircraft was subjected to large random initial state errors and to randomly directed crosswinds. The system was also tested for sensitivity to changes in the parameters of the aircraft and of the atmosphere. Performance of the optimal controller on all the three degree approaches was very good, and the control system proved to be reasonably insensitive to parametric uncertainties.
The analysis and preliminary design of an advanced technology transport aircraft flight control system using avionics and flight control concepts appropriate to the 1980-1985 time period are discussed. Specifically, the techniques and requirements of the flight control system were established, a number of candidate configurations were defined, and an evaluation of these configurations was performed to establish a recommended approach. Candidate configurations based on redundant integration of various sensor types, computational methods, servo actuator arrangements and data-transfer techniques were defined to the functional module and piece-part level. Life-cycle costs, for the flight control configurations, as determined in an operational environment model for 200 aircraft over a 15-year service life, were the basis of the optimum configuration selection tradeoff. The recommended system concept is a quad digital computer configuration utilizing a small microprocessor for input/output control, a hexad skewed set of conventional sensors for body rate and body acceleration, and triple integrated actuators.
Situation and command head-up displays were evaluated for both standard and two segment noise abatement night visual approaches in a fixed base simulation of a DC-8 transport aircraft. The situation display provided glide slope and pitch attitude information. The command display provided glide slope information and flight path commands to capture a 3 deg glide slope. Landing approaches were flown in both zero wind and wind shear conditions. For both standard and noise abatement approaches, the situation display provided greater glidepath accuracy in the initial phase of the landing approaches, whereas the command display was more effective in the final approach phase. Glidepath accuracy was greater for the standard approaches than for the noise abatement approaches in all phases of the landing approach. Most of the pilots preferred the command display and the standard approach. Substantial agreement was found between each pilot's judgment of his performance and his actual performance.
Conceptual design studies of pure and compound helicopter commercial short-haul transport aircraft for initial fabrication in 1980 were performed to determine their technical and economic feasibility. One-hundred-passenger configurations were optimized for minimum direct operating cost consistent with producibility and marketability, with emphasis on proper account of mass properties, performance and handling qualities adequacy, and suppression of internal and external noise. The effect of external noise constraints was assessed, in terms of gross weight and direct operating cost, for each aircraft.
Results of conceptual design studies of commercial rotary wing transport aircraft for the 1985 time period are presented. Two aircraft configurations, a tandem helicopter and a tilt rotor, were designed for a 200 nautical mile short haul mission with an upper limit of 100 passengers. In addition to the baseline aircraft two further designs of each configuration are included to assess the impact of external noise design criteria on the aircraft size, weight, and cost.
Bolted specimens representative of both upper and lower wing surface splices of a transport aircraft were designed and manufactured for static and random load tension and compression fatigue testing including ground-air-ground load reversals. The specimens were fabricated with graphite-epoxy composite material. Multiple tests were conducted at various load levels and the results were used as input to a statistical wearout model. The statically designed specimens performed very well under highly magnified fatigue loadings. Two large panels, one tension and compression, were fabricated for testing by NASA-LRC.
Data pertaining to nonmotion factors affecting ride quality of transport aircraft were obtained as part of NASA in-house and sponsored research studies carried out onboard commuter-airline and research aircraft. From these data, quantitative effects on passenger discomfort of seat width, seat legroom, change in cabin pressure, and cabin noise are presented. Visual cue effects are also discussed.
It is important to know whether the relatively high takeoff speeds of supersonic transport aircraft will change the exhaust noise levels of nozzles from those measured at static conditions. To gain some insight into this question, a modified F-106B aircraft was used to conduct flyover and static tests on a wedge nozzle. Flight velocity had an adverse effect on exhaust noise when compared with static results at the same relative jet velocity but a beneficial effect when compared with static results at the same absolute jet velocity. The wedge nozzle, which has a two-dimensional wedge surface rather than an axisymmetric plug surface, had a higher peak flyover noise level than the plug nozzle.
The technology and data base necessary for sound technical decisions regarding long haul supersonic cruise aircraft transportation systems are considered. The objectives and status of the research elements in the structures and materials program phase of the program are reviewed. Emphasis is placed on reductions in structural mass by research on advanced structural concepts, light-weight materials, improved loads, aeroelastic predictive techniques, and by development of efficient structural design procedures.
In theory, antisymmetric arrangements of wings and bodies can have smaller wave drag than corresponding mirror-symmetric arrangements. Thus, a long narrow oblique wing which presents the same aspect for two opposite directions of flight is potentially more efficient than corresponding (i.e., structurally equivalent) swept wing. The single continuous wing panel also adapts itself more readily to varying angles of obliquity, and hence, to varying flight speeds. Previous work on the aerodynamics and flight stability of oblique wing combinations is reviewed and a possible mode of application to transport aircraft operating at moderate supersonic speeds is suggested.
A study was made of the feasibility of a room-temperature scheme of accelerating fatigue tests for Mach 3 advanced supersonic transport aircraft. The test scheme used equivalent room-temperature cycles calculated for supersonic flight conditions. Verification tests were conducted using specimens representing titanium wing lower surface structure. Test-acceleration parameters were developed for the test with an auxiliary test set. Five specimens were tested with a flight-by-flight load and temperature spectrum to simulate typical Mach 3 operation. Two additional sets of five specimens were tested at room temperature to evaluate the test-acceleration scheme. The fatigue behavior of the specimens generally correlated well with the proposed correction method.
The aerodynamic design for a half-scale fan vehicle, which would have application on an advanced transport aircraft, is described. The single stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec 11,650 ft/sec). The fan and booster components are designed in a scale model flow size convenient for testing with existing facility and vehicle hardware. The design corrected flow per unit annulus area at the fan face is 215 kg/sec sq m (44.0 lb m/sec sq ft) with a hub-tip ratio of 0.38 at the leading edge of the fan rotor. This results in an inlet corrected airflow of 117.9 kg/sec (259.9 lb m/sec) for the selected rotor tip diameter if 90.37 cm (35.58 in.). The variable geometry inlet is designed utilizing a combination of high throat Mach number and acoustic treatment in the inlet diffuser for noise suppression (hybrid inlet). A variable fan exhaust nozzle was assumed in conjunction with the variable inlet throat area to limit the required area change of the inlet throat at approach and hence limit the overall diffusion and inlet length. The fan exit duct design was primarily influenced by acoustic requirements, including length of suppressor wall treatment; length, thickness and position on a duct splitter for additional suppressor treatment; and duct surface Mach numbers.
The structural design for a half-scale fan vehicle, which would have application on an advanced transport aircraft, is described. The single stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec (1,650 ft/sec). This mechanical design report describes the fan rotor design and the design of various structures of the vehicle; eg, stators, casings, splitters, seals, adapters, etc.
The acoustic design for a half-scale fan vehicle, which would have application on an advanced transport aircraft, is described. The single stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec (1,650 ft/sec). The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise is accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels (20 EPNdB) below current Federal Air Regulation noise standards for a full-scale fan at the takeoff, cutback, and approach conditions. Predicted unsuppressed and suppressed fore and aft maximum perceived noise levels indicate that the cutback condition is the most critical with respect to the goal, which is probably unattainable for that condition. This is also true for aft radiated noise in the approach condition.
The acoustic tests and data analysis for a 0.508-scale fan vehicle of a 111,300 newton (25,000 pound) thrust, full-size engine, which would have application on an advanced transport aircraft, is described. The single-stage advanced technology fan was designed to a pressure ratio of 1.8 at a tip speed of 503 m/sec (1,650 ft/sec) to achieve the desired pressure ratio in a single-stage fan with low radius ratio (0.38), and to maintain adequate stall margin. The fan has 44 tip-shrouded rotor blades and 90 outlet guide vanes. The two basic approaches taken in the acoustic design were: (1) minimization of noise at the source, and (2) suppression of the generated noise in the inlet and bypass exhaust duct. Suppression of the generated noise was accomplished in the inlet through use of the hybrid concept (wall acoustic treatment plus airflow acceleration suppression) and in the exhaust duct with extensive acoustic treatment including a splitter. The goal of the design was attainment of twenty effective perceived noise decibels (20 EPNdB) below current Federal Air Regulation noise standards for a full-scale fan at the takeoff, cutback, and approach conditions. The suppression goal of FAR 36-20 was not reached, but improvements in the technology of both front and aft fan-noise suppression were realized. The suppressed fan noise was shown to be consistent with the proposed federal regulation on aircraft noise.