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All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Supersonic airplane study and design

A supersonic airplane creates shocks which coalesce and form a classical N-wave on the ground, forming a double bang noise termed sonic boom. A recent supersonic commercial transport (the Concorde) has a loud sonic boom (over 100 PLdB) and low aerodynamic performance (cruise lift-drag ratio 7). To enhance the U.S. market share in supersonic transport, an airframer's market risk for a low-boom airplane has to be reduced. Computational fluid dynamics (CFD) is used to design airplanes to meet the dual constraints of low sonic boom and high aerodynamic performance. During the past year, a research effort was focused on three main topics. The first was to use the existing design tools, developed in past years, to design one of the low-boom wind-tunnel configurations (Ames Model 3) for testing at Ames Research Center in April 1993. The second was to use a Navier-Stokes code (Overflow) to support the Oblique-All-Wing (OAW) study at Ames. The third was to study an optimization technique applied on a Haack-Adams body to reduce aerodynamic drag.

Cheung, Samson↗
NASA NTRS1992

NASA F-16XL supersonic laminar flow control program overview

The viewgraphs and discussion of the NASA supersonic laminar flow control program are provided. Successful application of laminar flow control to a High Speed Civil Transport (HSCT) offers significant benefits in reductions of take-off gross weight, mission fuel burn, cruise drag, structural temperatures, engine size, emissions, and sonic boom. The ultimate economic success of the proposed HSCT may depend on the successful adaption of laminar flow control, which offers the single most significant potential improvements in lift drag ratio (L/D) of all the aerodynamic technologies under consideration. The F-16XL Supersonic Laminar Flow Control (SLFC) Experiment was conceived based on the encouraging results of in-house and NASA supported industry studies to determine if laminar flow control is feasible for the HSCT. The primary objective is to achieve extensive laminar flow (50-60 percent chord) on a highly swept supersonic wing. Data obtained from the flight test will be used to validate existing Euler and Navier Stokes aerodynamic codes and transition prediction boundary layer stability codes. These validated codes and developed design methodology will be delivered to industry for their use in designing supersonic laminar flow control wings. Results from this experiment will establish preliminary suction system design criteria enabling industry to better size the suction system and develop improved estimates of system weight, fuel volume loss due to wing ducting, turbocompressor power requirements, etc. so that benefits and penalties can be more accurately assessed.

Fischer, Michael C.↗
NASA NTRS1993

Experimental aerodynamic characteristics of a generic hypersonic accelerator configuration at Mach numbers 1.5 and 2.0

An experimental investigation of the static longitudinal and lateral-directional aerodynamic characteristics of a generic hypersonic research vehicle was conducted in the Langley Unitary Plan Wind Tunnel (UPWT). A parametric study was performed to determine the interference effects of various model components. Configuration variables included delta and trapezoidal canards; large and small centerline-mounted vertical tails, along with a set of wing-mounted vertical tails; and a set of model noses with different degrees of bluntness. Wing position was varied by changing the longitudinal location and the incidence angle. The test Mach numbers were 1.5 and 2.0 at Reynolds numbers of 1 x 10(exp 6) per foot, 2 x 10(exp 6) per foot, and 4 x 10(exp 6) per foot. Angle of attack was varied from -4 degrees to 27 degrees, and sideslip angle was varied from -8 degrees to 8 degrees. Generally, the effect of Reynolds number did not deviate from conventional trends. The longitudinal stability and lift-curve slope decreased with increasing Mach number. As the wing was shifted rearward, the lift-curve slope decreased and the longitudinal stability increased. Also, the wing-mounted vertical tails resulted in a more longitudinally stable configuration. In general, the lift-drag ratio was not significantly affected by vertical-tail arrangement. The best lateral-directional stability was achieved with the large centerline-mounted tail, although the wing-mounted vertical tails exhibited the most favorable characteristics at the higher angles of attack.

Walker, Ira J.↗
NASA NTRS1997

Neural Network Prediction of New Aircraft Design Coefficients

This paper discusses a neural network tool for more effective aircraft design evaluations during wind tunnel tests. Using a hybrid neural network optimization method, we have produced fast and reliable predictions of aerodynamical coefficients, found optimal flap settings, and flap schedules. For validation, the tool was tested on a 55% scale model of the USAF/NASA Subsonic High Alpha Research Concept aircraft (SHARC). Four different networks were trained to predict coefficients of lift, drag, moment of inertia, and lift drag ratio (C(sub L), C(sub D), C(sub M), and L/D) from angle of attack and flap settings. The latter network was then used to determine an overall optimal flap setting and for finding optimal flap schedules.

Norgaard, Magnus↗
NASA NTRS1960

A Flight Study of a Power-Off Landing Technique Applicable to Re-Entry Vehicles

A power-off landing technique, applicable to aircraft of configurations presently being considered for manned re-entry vehicles, has been developed and flight tested at Ames Research Center. The flight tests used two configurations of an airplane for which the values of maximum lift-drag ratios were 4.0 and 2.8. Twenty-four idle-power approaches were made to an 8000-foot runway with touchdown point and airspeed accuracies of +/-600 feet and +/-10 knots, respectively. The landing pattern used was designed to provide an explicitly defined flight path for the pilot and, yet, to require no external guidance other than the pilot's view from the cockpit. The initial phase of the approach pattern is a constant high-speed descent from altitude aimed at a ground reference point short of the runway threshold. At a specified altitude and speed, a constant g pull-out is made to a shallow flight path along which the air-plane decelerates to the touchdown point. Repeatability and safety are inherent because of the reduced number of variables requiring pilot judgment, and because of the fact that a missed approach is evident at speeds and altitudes suitable for safe ejection. The accuracy and repeatability of the pattern are indicated by the measured results. The proposed pattern appears to be particularly suitable for configurations having unusual drag variations with speed in the lower speed regime, since the pilot is not required to control speed in the latter portions of the pattern.

Bray, Richard S.↗
NASA NTRS1960

Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight

Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.

Wornom, Dewey E.↗
NASA NTRS1961

Wind-Tunnel Investigation of Paraglider Models at Supersonic Speeds

An investigation was made in the Langley Unitary Plan wind tunnel to determine the behavior of paraglider models at moderate to high supersonic speeds. The models were deployed from a sting in the supersonic stream and steady-state aerodynamic performance data were obtained. Maximum values of the lift-drag ratio were about 1.4 at a Mach number of 2.65 and about 1.2 at a Mach number of 4.65. The angles of attack over which the models could be flown were limited by unsteady behavior of the canopy.

Taylor, Robert T.↗
NASA NTRS1960

Hydrodynamic and Aerodynamic Characteristics of a Model of a Supersonic Multijet Water-Based Aircraft Equipped with Supercavitating Hydrofoils

The hydrodynamic and aerodynamic characteristics of a model of a multijet water-based Mach 2.0 aircraft equipped with hydrofoils have been determined. Takeoff stability and spray characteristics were very good, and sufficient excess thrust was available for takeoff in approximately 32 seconds and 4,700 feet at a gross weight of 225,000 pounds. Longitudinal and lateral stability during smooth-water landings were good. Lateral stability was good during rough-water landings, but forward location of the hydrofoils or added pitch damping was required to prevent diving. Hydrofoils were found to increase the aerodynamic lift-curve slope and to increase the aerodynamic drag coefficient in the transonic speed range, and the maximum lift-drag ratio decreased from 7.6 to 7.2 at the cruise Mach number of 0.9. The hydrofoils provided an increment of positive pitching moment over the Mach number range of the tests (0.6 to 1.42) and reduced the effective dihedral and directional stability.

McKann, Robert E.↗
NASA NTRS1959

Low-Speed Static Stability and Control Characteristics of a Model of a Right Triangular Pyramid Reentry Configuration

An investigation of the low-speed static stability and control characteristics of a model of a right triangular pyramid reentry configuration has been made in the Langley free-flight tunnel. The investigation showed that the model had generally satisfactory longitudinal and lateral static stability characteristics. The maximum lift-drag ratio was increased from about 3 to 5 by boattailing the base of the model.

Paulson, John W.↗
NASA NTRS1959

Launch, Low-Speed, and Landing Characteristics Determined from the First Flight of the North American X-15 Research Airplane

The first flight of the North American X-15 research airplane was made on June 8, 1959. This was accomplished after completion of a series of captive flights with the X-15 attached to the B-52 carrier airplane to demonstrate the aerodynamic and systems compatibility of the X-15//B-52 combination and the X-15 subsystem operation. This flight was planned as a glide flight so that the pilot need not be concerned with the propulsion system. Discussions of the launch, low-speed maneuvering, and landing characteristics are presented, and the results are compared with predictions from preflight studies. The launch characteristics were generally satisfactory, and the X-15 vertical tail adequately cleared the B-52 wing cutout. The actual landing pattern and landing characteristics compared favorably with predictions, and the recommended landing technique of lowering the flaps and landing gear at a low altitude appears to be a satisfactory method of landing the X-15 airplane. There was a quantitative correlation between flight-measured and predicted lift-drag-ratio characteristics in the clean configuration and a qualitative correlation in the landing configuration. A longitudinal-controllability problem, which became severe in the landing configuration, was evident throughout the flight and, apparently, was aggravated by the sensitivity of the side-located control stick. In the low-to-moderate angle-of-attack range covered, the longitudinal and directional stability were indicated to be adequate.

Finch, Thomas W.↗
NASA NTRS1954

Effect of a Leading-edge Flap upon the Lift, Drag, and Pitching Moment of an Airplane Employing a Thin, Unswept Wing

The effects of deflecting full-span, constant-chord, leading-edge flaps, having either round or sharp leading edges, upon the lift, drag,. and pitching moment characteristics of a model of an interceptor-type aircraft have been determined experimentally at subsonic and supersonic speeds. Results indicate that the variations of lift with angle of attack and of pitching moment with lift were unaffected by either the shape of the flap leading edge or flap deflection. Deflection of the flaps having either a round or sharp leading edge increased the drag at zero lift at both subsonic and supersonic speeds. In spite of the increase in the drag at zero lift, however, deflection of the flaps increased the maximum lift-drag ratio at subsonic speeds and had no deleterious effect at supersonic speeds.

Heitmeyer, John C.↗
NASA NTRS2012

NASA Technologies that Benefit Society

Applications developed on Earth of technology needed for space flight have produced thousands of spinoffs that contribute to improving national security, the economy, productivity and lifestyle. Over the course of it s history, NASA has nurtured partnerships with the private sector to facilitate the transfer of NASA-developed technology. For every dollar spent on research and development in the space program, it receives back $7 back in the form of corporate and personal income taxes from increased jobs and economic growth. A new technology, known as Liquid-metal alloy, is the result of a project funded by NASA s Jet Propulsion Lab. The unique technology is a blend of titanium, zirconium, nickel, copper and beryllium that achieves a strength greater than titanium. NASA plans to use this metal in the construction of a drill that will help for the search of water beneath the surface of Mars. Many other applications include opportunities in aerospace, defense, military, automotive, medical instrumentation and sporting goods.Developed in the 1980 s, the original Sun Tigers Inc sunlight-filtering lens has withstood the test of time. This technology was first reported in 1987 by NASA s JPL. Two scientists from JPL were later tasked with studying the harmful effects of radiation produced during laser and welding work. They came up with a transparent welding curtain that absorbs, filters and scatters light to maximize protection of human eyes. The two scientists then began doing business as Eagle Eye Optics. Each pair of sunglasses comes complete with ultraviolet protection, dual layer scratch resistant coating, polarized filters for maximum protection against glare and high visual clarity. Sufficient evidence shows that damage to the eye, especially to the retina, starts much earlier than most people realize. Sun filtering sunglasses are important. Winglets seen at the tips of airplane wings are among aviations most visible fuel-saving, performance enhancing technology. Aerodynamics centers on two majors forces-lift and drag. Lift enables the plane to fly and drag is the resistance encountered while moving through the air. The air coming off the end of a standard wing, in a vortex, causes significant drag and turbulence, forcing the spacing between takeoffs. A Langley engineer for NASA published finding in 1976 and verified by test in 1977 that winglets produce a 7 percent increase in lift-drag ratio with a 20 percent decrease in drag. This configuration reduces emissions, allows for greater range and carry more payload and the planes fly more quietly. Winglet technology has saved 2 billion gallons of jet fuel worldwide in 2010. Aerogel represents what technology experts believe to be the best insulation material ever invented. It is very light, flexible and can withstand temperatures of minus 3000 deg F. Aerogel products will be found in everything from clothing, to building insulation to space vehicles. Corpo Nove incorporated the Spaceloft version of the NASA-developed aerogel material into this jacket which was test during an Antarctic expedition.

Griffin, Amanda↗
NASA NTRS1991

Aerodynamic characteristics of scissor-wing geometries

A scissor-wing configuration, consisting of two independently sweeping-wing surfaces, is compared with an equivalent fixed-wing geometry baseline over a wide Mach number range. The scissor-wing configuration is shown to have a higher total lift-to-drag ratio than the baseline in the subsonic region primarily due to the slightly higher aspect ratio of the unswept scissor wing. In the transonic region, the scissor wing is shown to have a higher lift-to-drag ratio than the baseline for values of lift coefficient greater than 0.35. It is also shown that, through the use of wing decalage, the lift of the two independent scissor wings can be equalized. In the supersonic regime, the zero lift wave drag of the scissor-wing at maximum sweep is shown to be 50 and 28 percent less than the zero lift wave drag of the baseline at Mach numbers 1.5 and 3.0, respectively. In addition, a pivot-wing configuration is introduced and compared with the scissor wing. The pivot-wing configuration is shown to have a slightly higher total lift-to-drag ratio than the scissor wing in the supersonic region due to the decreased zero lift wave drag of the pivot-wing configuration.

Selberg, Bruce P.↗
NASA NTRS1984

Performance evaluation of the atmospheric phase of aeromaneuvering orbital transfer vehicles

Studies are underway to design reusable orbital transfer vehicles that would be used to transfer payloads from low-earth orbit to higher orbits and return. One promising concept is to use an atmospheric pass on the return leg to reduce the amount of fuel for the mission. This paper discusses a six-degree-of-freedom simulation analysis for two configurations, a low-lift-to-drag ratio configuration and a medium-lift-to-drag ratio configuration using both a predictive guidance technique and an adaptive guidance technique. Both guidance schemes were evaluated using the 1962 standard atmosphere and three atmospheres that had been derived from three entries of the Space Shuttle. The predictive technique requires less reaction control system activity for both configurations, but because of the limited number of updates and because each update used the 1962 standard atmosphere, the adaptive technique produces more accurate exit conditions.

Powell, R. W.↗
NASA NTRS2010

Orion Entry Performance-Based Center-of-Gravity Box

The Orion capsule has many performance requirements for its atmospheric entry trajectory. Requirements on landing accuracy, maximum heating rate, total heat load, propellant usage, and sensed acceleration must all be satised. It is desired to define a methodology to translate the many performance requirements for an atmospheric entry trajectory into language easily understood by vehicle designers in terms of an allowable center-of-gravity box. This is possible by noting that most entry performance parameters for a capsule vehicle are mainly determined by the lift-to-drag ratio of the vehicle. However, the lift-to- drag ratio should be considered a probabilistic quantity rather than deterministic, where variations in the lift-to-drag are caused by both aerodynamic and center-of-gravity un- certainties. This paper discusses the technique used by the Orion program to define the allowable dispersions in center-of-gravity to achieve the desired entry performance while accounting for aerodynamic uncertainty.

Rea, Jeremy R.↗
NASA NTRS2022

Computational Analysis of the X-57 Maxwell Airplane at Unpowered Conditions (Preliminary Fuselage)

The X-57 Maxwell is an all-electric airplane that implements a distributed electric propulsion system to demonstrate that high-efficiency electric propulsion can be integrated with aerodynamics to increase the performance of an airplane. To this end, distributed electric fans were installed on the wing to provide increased flow over the wing at the low takeoff and landing speeds of the X-57. The low-speed lift augmentation allows for a reduction in wing area for cruise optimization. The X-57 wing area was reduced to 42 percent of the wing area of the baseline aircraft, a Tecnam P2006T. With this reduced wing area and the electric propulsion system, it is estimated that the X-57 will cruise on less than one-third the total energy compared to the baseline aircraft. To meet the cruise performance goal at a Mach number of 0.233 at an altitude of 8000 feet, the X-57 has a cruise lift coefficient of 0.7516 and needs to have a cruise drag coefficient of 0.05423 or less. The USM3D computational solver was used to investigate the X-57 performance, without the distributed electric propulsion high-lift system operating. The unpowered X-57 performance is of interest to quantify if the X-57 can meet the cruise drag performance goal, and to document the lift performance of the very small wing at takeoff and landing conditions. The primary configurations investigated in this paper include the cruise configuration with no flap deflection, a takeoff configuration with a 10◦ flap deflection, and a landing configuration with a 30◦ flap deflection. The conditions for the cruise configuration were a flight unit Reynolds number of 1.32E+06 per foot, an altitude of 8000 feet, a Mach number of 0.233, and angles of attack from −2° to 24° . At the cruise lift coefficient of 0.7516, the computed drag coefficient is 0.05275. This computed drag is less than the drag coefficient of 0.05423 that is required to meet the X-57 airplane performance goal. However, the computational airplane is a completely smooth geometry and does not account for protuberance drag, nor the drag from steps and gaps in the actual X-57 airplane. Therefore, based upon the CFD drag calculation there is a 10-percent margin to account for some of the differences between the as-built metal fuselage and empennage construction, and the smooth computational geometry. The computed cruise drag also does not account for an induced drag reduction due to the wing-tip propellers and a drag reduction due to laminar flow achieved on the wing. The computed lift to drag ratio is 14.14 at the cruise lift coefficient of 0.7516, and the maximum computed lift to drag ratio is 15.8. The maximum lift coefficient for the cruise configuration was 2.13 at an angle of attack of 15°. The conditions for the takeoff configuration with a 10° flap deflection were a flight unit Reynolds number of 0.986E+06 per foot, an altitude of 2500 feet, a Mach number of 0.149, and angles of attack from −2° to 22°. The maximum lift coefficient for the takeoff configuration was 2.21 at an angle of attack of 16°. The conditions for the landing configuration with a 30° flap deflection were a flight unit Reynolds number of 0.922E+06 per foot, an altitude of 2500 feet, a Mach number of 0.139, and angles of attack from −2° to 24°. The maximum lift coefficient for the landing configuration was 2.58 and occurred at two angles of attack, 10° and 14°. Based on the unpowered maximum lift coefficient of 2.58 for the 30° flap deflection, along with computations of the distributed electric propulsion lift augmentation (not shown in this paper), the X-57 Maxwell is estimated to meet its powered landing goal of a maximum lift coefficient of 4.0.

X-57 Maxwell Airplane↗
NASA NTRS1990

Gamma guidance of trajectories for coplanar, aeroassisted orbital transfer

The optimization and guidance of trajectories for coplaner, aeroassisted orbital transfer (AOT) from high Earth orbit (HEO) to low Earth orbit (LEO) are examined. In particular, HEO can be a geosynchronous Earth orbit (GEO). It is assumed that the initial and final orbits are circular, that the gravitational field is central and is governed by the inverse square law, and that at most three impulses are employed: one at HEO exit, one at atmospheric exit, and one at LEO entry. It is also assumed that, during the atmospheric pass, the trajectory is controlled via the lift coefficient. The presence of upper and lower bounds on the lift coefficient is considered. First, optimal trajectories are computed by minimizing the total velocity impulse (hence, the propellant consumption) required for AOT transfer. The sequential gradient-restoration algorithm (SGRA) is used for optimal control problems. The optimal trajectory is shown to include two branches: a relatively short descending flight branch (branch 1) and a long ascending flight branch (branch 2). Next, attention is focused on guidance trajectories capable of approximating the optimal trajectories in real time, while retaining the essential characteristics of simplicity, ease of implementation, and reliability. For the atmospheric pass, a feedback control scheme is employed and the lift coefficient is adjusted according to a two-stage gamma guidance law. Further improvements are possible via a modified gamma guidance which is more stable with respect to dispersion effects arising from navigation errors, variations of the atmospheric density, and uncertainties in the aerodynamic coefficients than gamma guidance trajectory. A byproduct of the studies on dispersion effects is the following design concept. For coplaner aeroassisted orbital transfer, the lift-range-to-weight ratio appears to play a more important role than the lift-to-drag ratio. This is because the lift-range-to-weight ratio controls mainly the minimum altitude (hence, the peak heating rate) of the guidance trajectory; on the other hand, the lift-to-drag ratio controls mainly the duration of the atmospheric pass of the guidance trajectory.

Miele, A.↗
NASA NTRS2006

Computations of Viking Lander Capsule Hypersonic Aerodynamics with Comparisons to Ground and Flight Data

Comparisons are made between the LAURA Navier-Stokes code and Viking Lander Capsule hypersonic aerodynamics data from ground and flight measurements. Wind tunnel data are available for a 3.48 percent scale model at Mach 6 and a 2.75 percent scale model at Mach 10.35, both under perfect gas air conditions. Viking Lander 1 aerodynamics flight data also exist from on-board instrumentation for velocities between 2900 and 4400 m/sec (Mach 14 to 23.3). LAURA flowfield solutions are obtained for the geometry as tested or flown, including sting effects at tunnel conditions and finite-rate chemistry effects in flight. Using the flight vehicle center-of-gravity location (trim angle approx. equals -11.1 deg), the computed trim angle at tunnel conditions is within 0.31 degrees of the angle derived from Mach 6 data and 0.13 degrees from the Mach 10.35 trim angle. LAURA Mach 6 trim lift and drag force coefficients are within 2 percent of measured data, and computed trim lift-to-drag ratio is within 4 percent of the data. Computed trim lift and drag force coefficients at Mach 10.35 are within 5 percent and 3 percent, respectively, of wind tunnel data. Computed trim lift-to-drag ratio is within 2 percent of the Mach 10.35 data. Using the nominal density profile and center-of-gravity location, LAURA trim angle at flight conditions is within 0.5 degrees of the total angle measured from on-board instrumentation. LAURA trim lift and drag force coefficients at flight conditions are within 7 and 5 percent, respectively, of the flight data. Computed trim lift-to-drag ratio is within 4 percent of the data. Computed aerodynamics sensitivities to center-of-gravity location, atmospheric density, and grid refinement are generally small. The results will enable a better estimate of aerodynamics uncertainties for future Mars entry vehicles where non-zero angle-of-attack is required.

Edquist, Karl T.↗