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Dynamic stability of jet transport configuration with high thrust-weight ratio and externally blown jet flap
Wind tunnel investigation of jet transport airplane configuration with high thrust-weight ratio and external flow jet flap - graphs
Technology integration studies were made to examine the impact of emerging technologies on fighter aircraft. The technologies examined included advances in aerodynamics, controls, structures, propulsion, and systems and were those which appeared capable of being ready for application by the turn of the century. A primary impetus behind large increases in figher capability will be the rapid increase in fighter engine thrust-to-weight ratio. High thrust-weight engines, integrated with other advanced and emerging technologies, can result in small extremely maneuverable fighter aircraft that have thrust-weight ratios of 1.4+ and weight one-half as much as today's fighters. Future fighter aircraft requirements are likely to include a turn capability in excess of 7g's throughout much of the maneuver envelope, post-stall maneuverability, STOVL or VTOL, and a single engine for low cost.
A study of the principal flight parameters at booster separation was conducted to find the effect of each on the weight of the payload boosted into an earth orbit along a zero drag gravity turn trajectory. The parameters considered include Mach number (3 to 9), flight-path angle (10 deg to 55 deg), altitude (90,000 and 350,000 ft), inert weight ratio (0.05 to 0.15), and thrust-weight ratio (1.5 to 2.5), with a specific impulse of 289 seconds. Both transfer ellipse and direct orbit trajectories were considered. With either trajectory method, payload weight was highest for low initial flight-path angles and high initial Mach numbers. Of course, high initial Mach numbers require greater energy expenditures from the booster. Changes in initial altitude had little effect on payload weight, and only small gains were evident when thrust-weight ratio was increased.
A systems study was conducted on an aircraft concept, representative of a supersonic-cruise military aircraft (supercruiser). The study results indicate that supersonic ranges in excess of 4000 n.mi. at a Mach number of 2.62 are possible with a 500 lbf class aircraft. Trade studies, to determine the sensitivity of supersonic range to parameters which would improve maneuverability, indicate that thrust-weight ratios of as much as 0.5 can be used without significantly decreasing supersonic range; however, increasing the thrust-weight ratio to 1.0 decreases the range capability by about 1100 n.mi. The range penalty for increasing the aircraft limit load-factor from 4.0 to 9.0 is about 500 n.mi. The increased fuel volume of several configurations improved the supersonic range capability by about 1200 n.mi. but, due to associated losses in supersonic L/D, had an insignificant effect on the range at a Mach number of 2.62.
Effect of thrust-weight ratios and specific impulses on trajectory parameters for earth orbit hyperbolic escape and entry
A brief study of the effects of turbofan-engine bypass ratio on Breguet cruise range and take-off distance for subsonic cruise aircraft showed significant differences between short- and long-haul aircraft designs. Large thrust lapse rates at high bypass ratios caused severe reductions in cruise range for short-haul aircraft because of increases in propulsion system weight. Long-haul aircraft, with a higher fuel fraction (ratio of propulsion weight plus total fuel weight to gross take-off weight), are less sensitive to propulsion-system weight and, accordingly, were not significantly affected by bypass-ratio variations. Both types of aircraft have shorter take-off distances at higher bypass ratios because of higher take-off thrust-weight ratios.
A LOX/LH2 high/variable mixture ratio booster upper stage is described. The engine has high thrust-weight ratio as a booster and high specific impulse as an upper stage engine. Operation at high mixture ratio utilizes the propellants at high bulk density. The engine may use multiple turbopump-preburners for higher thrust ratings. The engine uses the full flow cycle to obtain minimum turbine inlet temperatures for a given chamber pressure and to avoid interpropellant shaft seals and other single point failure modes. A portion of the liquid hydrogen is used to regeneratively cool the thrust chamber assembly. The warmed hydrogen coolant is then used to drive the fuel boost turbopump. All propellants arrive at the gas-gas injector ready to burn. Shear mixing of the parallel flowing high velocity, low density fuel-rich gases with the high density, low velocity oxidizer-rich gases provides complete combustion with a modest chamber volume. Combustion stability is assured by the injection of the heated fuel-rich gases and the comparatively low volume ratio of the propellants before and after combustion. The high area ratio nozzle skirt is fitted with a low area ratio nozzle skirt insert for optimum low altitude performance. The overall engine characteristics make it a candidate for ALS, Shuttle-C, LRB, and SSTO applications.
Multiorbit injection earth departure technique for optimal thrust-weight ratio in manned interplanetary missions using NERVA engine
Ascent performance characteristics were studied for a vertical-take-off, horizontal-landing, single-stage-to-orbit transport vehicle powered by hydrogen fuel rocket with a mixture of fixed- and dual-position nozzles. The analysis was made by systematically varying two sets of trajectory similarity parameters based on the propulsive and aerodynamic characteristics of the vehicle and by calculating a trajectory for each combination of the parameters. The propulsion parameters were the initial thrust-weight ratio, engine combination, and the two expansion ratios of the dual-position rocket nozzles. The aerodynamic parameters were the ratio of reference area to initial weight and the ratio of maximum allowable normal force to initial weight. A first-order analysis was carried out to determine the effect on the performance of including the engine mass penalty. This analysis indicates that the configuration with the lowest initial mass for a given payload requires all dual-position nozzles with initial expansion ratio of 50 and a final expansion ratio of 150.
The results of a parametric study of ascent performance are presented for a vertical take off, horizontal landing, single stage earth-to-orbit transport vehicle. Two dual fueled concepts, series burn and parallel burn, were investigated, both of which utilized dual position rocket nozzles. The analysis was made by systematically varying a set of propulsion similarity parameters, initial thrust-weight ratio, the proportion of the thrust due to dual position nozzle engines, expansion ratios of the rocket nozzle, and the relative split between the two fuels, hydrogen and hydrocarbon. The data are presented as a series of curves of mass ratio plotted against each of the similarity parameters for various combinations of the other similarity parameters.
A procedure that allows rapid preliminary evaluations of the vertical, short, and normal takeoff performance of supersonic cruise aircraft concepts was developed into a numerical computer program. The program is used to determine the effects on takeoff performance of various parameters, such as thrust-weight ratio, wing loading, thrust vector angle, and flap setting. Ramp-assisted takeoffs for overloaded configurations typical of a ground-attack mission are included. The effects of wind on the takeoff performance are also considered.
A wind-tunnel investigation at low speeds has been made to study the aerodynamic characteristics of a small-scale sweptback-wing Jet-transport model equipped with an external-flow jet-augmented double slotted flap. Included in the investigation were tests of the wing alone to study the effects of varying the spanwise extent of blowing on the full-span flap. The results indicated that the double-slotted-flap arrangement of the present investigation was more efficient in terms of lift and drag than were the external-flow single-slotted-flap arrangements previously tested and gave a substantial reduction In the thrust-weight ratio required for a given lift coefficient under trimmed drag conditions. An increase in the spanwise extent of blowing on the full-span flap was also found to increase the efficiency of the model in terms of the lift and drag but, as would be expected on a sweptback-wing configuration, was accompanied by significant increases in negative pitching moment.
The take-off distances over a 35-foot obstacle have been determined for a supersonic transport configuration characterized by a low maximum lift coefficient at a high angle of attack and by high drag due to lift. These distances were determined analytically by means of an electronic digital computer. The effects of rotation speed, rotation angle, and rotation time were determined. A few configuration changes were made to determine the effects of thrust-weight ratio, wing loading, maximum lift coefficient, and induced drag on the take-off distance. The required runway lengths based on Special Civil Air Regulation No. SR-422B were determined for various values of rotation speed and compared with those based on full engine power. Increasing or decreasing the rotation speed as much as 5 knots from the value at which the minimum take-off distance occurred increased the distance only slightly more than 1 percent for the configuration studied. Under-rotation by 1 deg to 1.5 deg increased the take-off distance by 9 to 15 percent. Increasing the time required for rotation from 3 to 5 seconds had a rather small effect on the take-off distance when the values of rotation speed were near the values which result in the shortest take-off distance. When the runway length is based on full engine power rather than on SR-422B, the rotation speed which results in the shortest required runway length is 10 knots lower and the runway length is 4.3 percent less.
Fixed-cockpit piloted simulator studies of delta-planform and variable-wing-sweep supersonic transport configurations are being conducted at the Ames Research Center to investigate the handling qualities and certification requirements related to the take-off maneuver. Validation of the simulation was achieved by duplicating the take-off certification program of a subsonic jet transport. Evaluation of the simulator was made by NASA pilots as well as company and FAA pilots involved in the actual certification flights of the airplane. The present paper is limited to a discussion of normal take-off, minimum control speed (ground), rotation characteristics, and initial climbout. Comparisons of the take-off characteristics are made between the supersonic transport and the current class of subsonic jet transports. Results indicate that minimum control speed (ground) characteristics are a function of thrust-weight ratio, the time provided for SST rotation should be at least as long as that for the subsonic jet transports, abused take-offs are more likely to result in tail scrapes, and climbout below the minimum drag speed requires that the pilot carefully monitor airspeed.
Although the major NASA research effort is directed toward XB-70-2, which will not enter its flight program until the summer of 1965, a limited amount of information is available from the early flights of the XB-70-1 airplane. Initial take-off and landing performance data have generally substantiated predictions and indicate no unforeseen problems for this class of vehicle. Vertical velocities at impact are of the same order of magnitude as those being experienced by present-day subsonic jets. The XB-70 distances from brake release to lift-off graphically illustrate the advantage of the increased thrust-weight ratio of the supersonic cruise vehicle. The landing loads are well within the design limits up to the highest vertical velocities encountered to date, and recorded data show the response at the pilot station to be somewhat greater than that recorded at the center of gravity. Persistent shaking has been encountered in flight at subsonic speed. The cause of the excitation is not known at present but the oscillation does not appear to be conventional buffeting. The oscillation occurrence drops off appreciably at supersonic speeds and can be correlated with atmospheric turbulence. The stability and control characteristics at subsonic speeds appear satisfactory with stability augmentation on and off. A longitudinal trim discrepancy from predictions has been noted in the transonic region which appears to be decreasing with increasing supersonic speed. The supersonic handling qualities are considered adequate with stability augmentation off; however, sensitive lateral control has resulted in small pilot-induced oscillations.
Relationships between rocket weight, velocity, and thrust power - propulsive efficiency
Electric thrustor development reviewed with stress on lifetime, efficiency, and potential use in space program