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NASA NTRS1940

Aerodynamic characteristics of horizontal tail surfaces

Collected data are presented on the aerodynamic characteristics of 17 horizontal tail surfaces including several with balanced elevators and two with end plates. Curves are given for coefficients of normal force, drag, and elevator hinge moment. A limited analysis of the results has been made. The normal-force coefficients are in better agreement with the lifting-surface theory of Prandtl and Blenk for airfoils of low aspect ratio than with the usual lifting-line theory. Only partial agreement exists between the elevator hinge-moment coefficients and those predicted by Glauert's thin-airfoil theory.

Silverstein, Abe↗

NASA NTRS1948

Tank Tests of a Powered Model of a Compression Plane, NACA Model 171A-2

The compression plane is intended for operation on or close to the surface of the water, and has a hull with a concave bottom which forms the upper surface of a tunnel into which air is forced under pressure to support part of the load. The results of the tests made in Langley tank no. 1 include values of the horizontal forces, trimming moment, and static pressure in the tunnel for a wide range of loads and speeds and two power conditions, and are presented in the form of curves against speed with load as a parameter. The results are scaled up to 10 times the model size for three conditions at which the model is self-propelled at a steady speed. Lift is obtained from the static pressure of air in the tunnel. In general, the ratio of the gross load to the total resistance increases with increase in load and decrease in speed. This ratio varies between l-7 and 5.7 at high speeds and has a maximum value of 7. The total resistance is nearly the same for both power conditions except at low speeds and heavy loads. No abrupt change in forces on the hull or flow around the hull occurs in. the region of zero draft. The centers of pressure are generally far aft. At the most efficient trim (1.2'), considerable bow-up moment would be required for practicable operation. There is no abrupt transition from the air-borne to the water- borne condition.

Mottard, Elmo J.↗

NASA NTRS1971

Flight-determined aerodynamic stability and control derivatives of the M2-F2 lifting body vehicle at subsonic speeds

Aerodynamic derivatives were obtained for the M2-F2 lifting body flight vehicle in the subsonic flight region between Mach numbers of 0.41 and 0.64 and altitudes of 7000 feet to 45,000 feet. The derivatives were determined by a flight time history curve-fitting process utilizing a hybrid computer. The flight-determined derivatives are compared with wind-tunnel and predicted values. Modal-response characteristics, calculated from the flight derivatives, are presented.

Kempel, R. W.↗

NASA NTRS1971

Mars surface sample return missions via solar electric propulsion

The characteristics and capabilities are described of solar electric propulsion (SEP) for performing Mars Surface Sample Return (MSSR) missions. The scope of the study emphasizes trajectory/payload analysis and the comparison of mission/system tradeoff options. The MSSR mission is examined only for the 1981-82 launch opportunity. Several other study constraints which bear directly on the results obtained are: (1) return samples in the range 5-25 kg, (2) use of lifting (offset C.G.) atmospheric entry at Mars which allows a low ratio (1.25) of entry weight to landed weight, and (3) rendezvous and docking in Mars orbit. Major results of the study are presented as performance curves of earth departure mass versus sample size for a number of different mission/system options. These options represent a spectrum of trip time, launch vehicle capability, combinations of low-thrust and ballistic maneuvers, chemical retro type, and earth recovery mode.

Spadoni, D. J.↗

NASA NTRS1989

Performance of an aero-space plane propulsion nozzle

An inviscid and viscous analysis is provided for an exposed half nozzle that is used with a scramjet for thrust generation. The analysis is based on the inviscid theory of a two-dimensional, minimum length nozzle with a curved inlet surface, where the flow may be sonic or supersonic. Inlet conditions are prescribed and the gas is assumed to be perfect. Viscous, and when appropriate inviscid, nondimensional parametric results are provided for the thrust, lift, heat transfer, pitching moment, and a variety of boundary-layer thicknesses. In addition to global results, wall distributions of pressure, heat transfer, etc., are provided. The analysis demonstrates that the nozzle produces a considerable lift force whose magnitude may exceed the thrust and a significant pitching moment. The thrust is quite sensitive to the inlet Mach number; it rapidly decreases as the inlet Mach number increases There is little loss in the thrust as the nozzle's downstream wall is truncated. The corresponding decrease in lift and the pitching moment is moderate.

Emanuel, George↗

NASA NTRS1978

A three-dimensional potential-flow program with a geometry package for input data generation

Information needed to run a computer program for the calculation of the potential flow about arbitrary three dimensional lifting configurations is presented. The program contains a geometry package which greatly reduces the task of preparing the input data. Starting from a very sparse set of coordinate data, the program automatically augments and redistributes the coordinates, calculates curves of intersection between components, and redistributes coordinates in the regions adjacent to the intersection curves in a suitable manner for use in the potential flow calculations. A brief summary of the program capabilities and options is given, as well as detailed instructions for the data input, a suggested structure for the program overlay, and the output for two test cases.

Halsey, N. D.↗

NASA NTRS2001

The Impact on Simulated Storm Structure and Intensity of Variations in the Lifted Condensation Level and the Level of Free Convection

The sensitivities of convective storm structure and intensity to changes in the altitudes of the prestorm environmental lifted condensation level and level of free convection axe studied using a full-physics three-dimensional cloud model. Matrices of simulations are conducted for a range of LCL=LFC altitudes, using a single moderately-sheared curved hodograph trace in conjunction with 1 convective available potential energy values of either 800 or 2000 J/kg, with the matrices consisting of all four combinations of two distinct choices of buoyancy and shear profile shape. For each value of CAPE, the LCL=LFC altitudes are also allowed to vary in a series of simulations based on the most highly compressed buoyancy and shear profiles for that CAPE, with the environmental buoyancy profile shape, subcloud equivalent potential temperature, subcloud lapse rates of temperature and moisture, and wind profile held fixed. For each CAPE, one final simulation is conducted using a near optimal LFC, but a lowered LCL, with a neutrally buoyant environmental thermal profile specified in between. Results show that, for the buoyancy-starved small-CAPE environments, the simulated storms are supercells and are generally largest and most intense when LCL=LFC altitudes lie in the approximate range 1.5-2.5 km above the surface. The simulations show similar trends for the shear-starved large-CAPE environments, except that conversion from supercell to multicell morphology frequently occurs when the LCL is high. For choices of LCL=LFC height within the optimal 1.5-2.5 km range, peak storm updraft overturning efficiency may approaches unity relative to parcel theory, while for lower LCL=LFC heights, overturning efficiency is reduced significantly. The enhancements of overturning efficiency and updraft diameter with increasing LFC height are shown to be the result of systematic increases in the mean equivalent potential temperature of the updraft at cloud base. For the shear-starved environments, the tendency for outflow dominance is eliminated, but a large overturning efficiency maintained, when a low LCL is used in conjunction with a high LFC. The result regarding outflow dominance at high LCL is consistent with expectations, but the beneficial effect of a high LFC on convective overturning efficiency has not previously been widely recognized. The simulation findings here also appear to be consistent with statistics from previous severe storm environment climatologies, but provide a new framework for interpreting those statistics.

McCaul, Eugene W., Jr.↗

NASA NTRS2010

Large Spun Formed Friction-Stir Welded Tank Domes for Liquid Propellant Tanks Made from AA2195: A Technology Demonstration for the Next Generation of Heavy Lift Launchers

Improving structural efficiency while reducing manufacturing costs are key objectives when making future heavy-lift launchers more performing and cost efficient. The main enabling technologies are the application of advanced high performance materials as well as cost effective manufacture processes. This paper presents the status and main results of a joint industrial research & development effort to demonstrate TRL 6 of a novel manufacturing process for large liquid propellant tanks for launcher applications. Using high strength aluminium-lithium alloy combined with the spin forming manufacturing technique, this development aims at thinner wall thickness and weight savings up to 25% as well as a significant reduction in manufacturing effort. In this program, the concave spin forming process is used to manufacture tank domes from a single flat plate. Applied to aluminium alloy, this process allows reaching the highest possible material strength status T8, eliminating numerous welding steps which are typically necessary to assemble tank domes from 3D-curved panels. To minimize raw material costs for large diameter tank domes for launchers, the dome blank has been composed from standard plates welded together prior to spin forming by friction stir welding. After welding, the dome blank is contoured in order to meet the required wall thickness distribution. For achieving a material state of T8, also in the welding seams, the applied spin forming process allows the required cold stretching of the 3D-curved dome, with a subsequent ageing in a furnace. This combined manufacturing process has been demonstrated up to TRL 6 for tank domes with a 5.4 m diameter. In this paper, the manufacturing process as well as test results are presented. Plans are shown how this process could be applied to future heavy-lift launch vehicles developments, also for larger dome diameters.

Stachulla, M.↗

NASA NTRS1974

Effect of canard location and size on canard-wing interference and aerodynamic center shift related to maneuvering aircraft at transonic speeds

A generalized wind-tunnel model, typical of highly maneuverable aircraft, was tested in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.70 to 1.20 to determine the effects of canard location and size on canard-wing interference effects and aerodynamic center shift at transonic speeds. The canards had exposed areas of 16.0 and 28.0 percent of the wing reference area and were located in the chord plane of the wing or in a position 18.5 percent of the wing mean geometric chord above or below the wing chord plane. Two different wing planforms were tested, one with leading-edge sweep of 60 deg and the other 44 deg; both wings had the same reference area and span. The results indicated that the largest benefits in lift and drag were obtained with the canard above the wing chord plane for both wings tested. The low canard configuration for the 60 deg swept wing proved to be more stable and produced a more linear pitching-moment curve than the high and coplanar canard configurations for the subsonic test Mach numbers.

Gloss, B. B.↗

NASA NTRS1973

Point and path performance of light aircraft: A review and analysis

The literature on methods for predicting the performance of light aircraft is reviewed. The methods discussed in the review extend from the classical instantaneous maximum or minimum technique to techniques for generating mathematically optimum flight paths. Classical point performance techniques are shown to be adequate in many cases but their accuracies are compromised by the need to use simple lift, drag, and thrust relations in order to get closed form solutions. Also the investigation of the effect of changes in weight, altitude, configuration, etc. involves many essentially repetitive calculations. Accordingly, computer programs are provided which can fit arbitrary drag polars and power curves with very high precision and which can then use the resulting fits to compute the performance under the assumption that the aircraft is not accelerating.

Smetana, F. O.↗

NASA NTRS1975

Interim prediction method for externally blown flap noise

An interim procedure for predicting externally blown flap (EBF) noise spectra anywhere below a powered lift aircraft is presented. Both engine-under-the-wing and engine-over-the-wing EBF systems are included. The method uses data correlations for the overall sound pressure level based on nozzle exit area and exhaust velocity along with OASPL directivity curves and normalized one-third-octave spectra. Aircraft motion effects are included by taking into account the relative motion of the source with respect to the observer and the relative velocity effects on source strength.

Dorsch, R. G.↗

NASA NTRS1986

High speed viscous flow calculations about complex configurations

Applications of the NASA Ames Parabolized Navier-Stokes (PNS) code to a variety of complex generic configurations is presented. The algorithm, boundary conditions, initial conditions, and grid generators are discussed as applied to these configurations. The PNS code was used as the mainline procedure to numerically simulate the viscous supersonic flow over these generic configurations. The turbulence model that was used in this study is the Baldwin-Lomax model. The boundary conditions are the usual viscous no slip at the wall, and a characteristic procedure is used to fit the bow shock wave which is the outermost boundary. An elliptic grid generator is employed to discretize the flow domain. In addition, an equilibrium air capability has been incorporated into the code. It uses the curve fits of Tannehill, et al. The flow regimes vary from a Mach number of 2 up to 25. Both laminar and turbulent flow are considered. Varying angles of attack have also been computed. Configurations vary from simple cone-type bodies to lifting winged bodies, such as the space shuttle or the generic supersonic cruise fighter.

Chaussee, D. S.↗

NASA NTRS1937

Theoretical span loading and moments of tapered wings produced by aileron deflection

The effect of tapered ailerons on linearly tapered wings is theoretically determined. Four different aileron spans are considered for each of three wing aspect ratios and each of four wing taper ratios. The change in lift on one half of the wing, the rolling moment, the additional induced drag, and the yawing moment, due to aileron deflection, are represented by non dimensional coefficients. Similar coefficients are given for the damping and yawing moments, the additional drag, and the change in lift, due to rolling. It was found possible to effect a fairly close agreement between the theoretical and experimental rolling moments by introducing into the theoretical expression for the rolling moment an effective change in angle of attack obtained from an analysis of flap data. The theoretical curves show that the highly tapered wing with long ailerons has a lower ratio of yawing to rolling moment and a lower additional induced drag than wings with less taper.

Pearson, H A↗

NASA NTRS1957

Interference Effects of Fuselage-Stored Missiles on Inlet Duct Model of an Interceptor-Type Aircraft at Mach Numbers 1.5 to 1.9

The effect of missile armament on the performance of an interceptor-type aircraft model has been determined at Mach numbers 1.5, 1.7, and 1.9 and at angles of attack to 19 deg. With this configuration missiles were carried in a bay located on the bottom of the aircraft fuselage and mounted to a rotatable missile door. Rotation of the door then brought the missiles into the external or firing position. The aircraft model was characterized by triangular-shaped normal-shock inlets located at the wing roots. Relatively short and curved subsonic diffusers fed simulated twin side-by-side turbojet engines. Inasmuch as the missile bay extended considerably ahead of the inlet station, rotation of the missile door created considerable disturbance of the flow entering the inlets. In comparison with the internal missile arrangement, the external missile configurations increased the model lift, drag, and pitching moment. While the diffuser-exit flow distortion and static-pressure fluctuations were not greatly affected, diffuser total-pressure recovery was reduced as much as 0.058 at Mach number 1.9 for one missile configuration. The most detrimental effect of missile-door rotation occurred at the transient door positions, or with the door halfwzy between the missiles-in and -out conditions. At this door position the flow into the inlets was highly asymmetrical. Although the performance of both left and right ducts was generally reduced, the inlet duct on the cavity side of the missile door was most severely penalized, becoming unstable recovery losses and increases in flow distortion. The installation of fuselage fences along the missile bay was only partially effective in reducing these losses.

Piercy, T. G.↗

NASA NTRS2017

Diagnostic Techniques to Elucidate the Aerodynamic Performance of Acoustic Liners

In support of Topic A.2.8 of NASA NRA NNH10ZEA001N, the University of Florida (UF) has investigated the use of flow field optical diagnostic and micromachined sensor-based techniques for assessing the wall shear stress on an acoustic liner. Stereoscopic particle image velocimetry (sPIV) was used to study the velocity field over a liner in the Grazing Flow Impedance Duct (GFID). The results indicate that the use of a control volume based method to determine the wall shear stress is prone to significant error. The skin friction over the liner as measured using velocity curve fitting techniques was shown to be locally reduced behind an orifice, relative to the hard wall case in a streamwise plane centered on the orifice. The capacitive wall shear stress sensor exhibited a linear response for a range of shear stresses over a hard wall. PIV over the liner is consistent with lifting of the near wall turbulent structure as it passes over an orifice, followed by a region of low wall shear stress.

June, Jason↗

NASA NTRS1997

Vortex leading edge flap assembly for supersonic airplanes

A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.

Rudolph, Peter K. C.↗

NASA NTRS2013

Radial Internal Material Handling System (RIMS) for Circular Habitat Volumes

On planetary surfaces, pressurized human habitable volumes will require a means to carry equipment around within the volume of the habitat, regardless of the partial gravity (Earth, Moon, Mars, etc.). On the NASA Habitat Demonstration Unit (HDU), a vertical cylindrical volume, it was determined that a variety of heavy items would need to be carried back and forth from deployed locations to the General Maintenance Work Station (GMWS) when in need of repair, and other equipment may need to be carried inside for repairs, such as rover parts and other external equipment. The vertical cylindrical volume of the HDU lent itself to a circular overhead track and hoist system that allows lifting of heavy objects from anywhere in the habitat to any other point in the habitat interior. In addition, the system is able to hand-off lifted items to other material handling systems through the side hatches, such as through an airlock. The overhead system consists of two concentric circle tracks that have a movable beam between them. The beam has a hoist carriage that can move back and forth on the beam. Therefore, the entire system acts like a bridge crane curved around to meet itself in a circle. The novelty of the system is in its configuration, and how it interfaces with the volume of the HDU habitat. Similar to how a bridge crane allows coverage for an entire rectangular volume, the RIMS system covers a circular volume. The RIMS system is the first generation of what may be applied to future planetary surface vertical cylinder habitats on the Moon or on Mars.

Howe, Alan S.↗

NASA NTRS1978

A geometry package for generation of input data for a three-dimensional potential-flow program

The preparation of geometric data for input to three-dimensional potential flow programs was automated and simplified by a geometry package incorporated into the NASA Langley version of the 3-D lifting potential flow program. Input to the computer program for the geometry package consists of a very sparse set of coordinate data, often with an order of magnitude of fewer points than required for the actual potential flow calculations. Isolated components, such as wings, fuselages, etc. are paneled automatically, using one of several possible element distribution algorithms. Curves of intersection between components are calculated, using a hybrid curve-fit/surface-fit approach. Intersecting components are repaneled so that adjacent elements on either side of the intersection curves line up in a satisfactory manner for the potential-flow calculations. Many cases may be run completely (from input, through the geometry package, and through the flow calculations) without interruption. Use of the package significantly reduces the time and expense involved in making three-dimensional potential flow calculations.

Halsey, N. D.↗