SEARCH · Search NASA
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Definitions of methods of measurements of supporting and control surface areas are presented. Methods for measuring the supporting surface, i.e., the wing area, and the control surfaces, i.e., the horizontal tail area, the vertical tail area, and the trailing control surface areas are defined. Illustrations of each of the areas are included.
Explore the source record for details and available documents.
The National Aeronautics and Space Administration has actively participated in the development of hot structures technology for application to hypersonic flight systems. Hot structures have been developed for vehicles including the X-43A, X-37, and the Space Shuttle. These trans-atmospheric and atmospheric entry flight systems that incorporate hot-structures technology are lighter weight and require less maintenance than those that incorporate parasitic, thermal-protection materials that attach to warm or cool substructure. The development of hot structures requires a thorough understanding of material performance in an extreme environment, boundary conditions and load interactions, structural joint performance, and thermal and mechanical performance of integrated structural systems that operate at temperatures ranging from 1500 C to 3000 C, depending on the application. This paper will present recent advances in the development of hot structures, including development of environmentally durable, high temperature leading edges and control surfaces, integrated thermal protection systems, and repair technologies. The X-43A Mach-10 vehicle utilized carbon/carbon (C/C) leading edges on the nose, horizontal control surface, and vertical tail. The nose and vertical and horizontal tail leading edges were fabricated out of a 3:1 biased, high thermal conductivity C/C. The leading edges were coated with a three-layer coating comprised of a SiC conversion of the C/C, followed by a CVD layer of SiC, followed by a thin CVD layer of HfC. Work has also been performed on the development of an integrated structure and was focused on both hot and warm (insulated) structures and integrated fuselage/tank/TPS systems. The objective was to develop integrated multifunctional airframe structures that eliminate fragile external thermal-protection systems and incorporate the insulating function within the structure. The approach taken to achieve this goal was to develop candidate hypersonic airframe concepts, including structural arrangement, load paths, thermal-structural wall design, thermal accommodation features, and integration of major components, optimize thermalstructural configurations, and validate concepts through a building block test program and generate data to improve and validate analytical and design tools.
A rapid shape parameterization tool called PROTEUS is developed for aircraft shape optimization. This tool can be applied directly to any aircraft geometry that has been defined in PLOT3D format, with the restriction that each aircraft component must be defined by only one data block. PROTEUS has eight types of parameterization schemes: planform, wing surface, twist, body surface, body scaling, body camber line, shifting/scaling, and linear morphing. These parametric schemes can be applied to two types of components: wing-type surfaces (e.g., wing, canard, horizontal tail, vertical tail, and pylon) and body-type surfaces (e.g., fuselage, pod, and nacelle). These schemes permit the easy setup of commonly used shape modification methods, and each customized parametric scheme can be applied to the same type of component for any configuration. This paper explains the mathematics for these parametric schemes and uses two supersonic configurations to demonstrate the application of these schemes.
Previous research on airplane spinning and recovery has shown that at potential spin conditions (high angles of attack with rotation) the horizontal tail, depending upon its location, can create a wake about the vertical tail and rudder which can adversely affect airplane spin and recovery characteristics. Many methods of altering the tail geometry to modify these interference effects were investigated for improving airplane spin and recovery characteristics. Examples of changes includes relocation of the horizontal tail, increasing control surface travel, and use of a 'flip tail' that can be rotated to extreme angles for spin recovery. A device is provided which improves the spin recovery characteristics of aircraft which involves attaching the horizontal tail of the aircraft to the aircraft such that a gap remains between the root end of each horizontal tail section and the fuselage or vertical tail of the aircraft. The gaps measure between about 15 and 30 percent of the tail semispan. The gaps may be covered by shields which are released should a spin occur.
Explore the source record for details and available documents.
For the stall proof airplane there should be a stabilizing pitching moment below the stall angle of attack of sufficient amount to prevent the attainment of the stall angle of attack which cannot be over-ridden by control deflections. This paper presents (1) a development of the moment equations to show the theoretical considerations of a stall proof airplane and (2) the nonlinear moment characteristics that must be obtained to satisfy the stall proof requirements. In addition, it is shown that an aerodynamic spoiler located on the under surface of the horizontal tail can be designed to meet these requirements of a stall proof airplane. Wind tunnel results are shown to validate assumptions and predictions.
The tests were conducted at Mach numbers from 0.40 to 0.90, at angles of attack up to 45 deg for the lower Mach numbers, and at angles of sideslip up to 15 deg. The model variations under study included adding a canard surface and deflecting horizontal tails, ailerons, and rudders.
An investigation was conducted in the Langley 16 Foot Transonic Tunnel to determine the lateral directional aerodynamic characteristics of a fully metric 0.04 scale model of the F-15 three surface configuration (canards, horizontal tails) with twin two dimensional nozzles and twin axisymmetric nozzles installed. The effects of two dimensional nozzle in flight thrust reversing and rudder deflection were also determined. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.20 over an angle of attack range from -2 deg to 15 deg. Reynolds number varied from 2.6 million to 3.8 million. Angle of sideslip was set at approximately 0 deg and -5 deg for all configurations and at -10 deg for selected configurations.
Pressure-distribution tests of an N.A.C.A. 0009 airfoil with a 50-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made in the N.A.C.A. 4- by 6- foot vertical tunnel. The tests supplied aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. The experimental results and values computed by analytical methods are in good agreement for small flap and tab deflections. The results of the tests indicated that the effectiveness of all three tab sizes in reducing flap hinge moments decreased with increasing flap deflection.
Pressure-distribution tests of an NACA 0009 airfoil with a 30-percent-chord plain flap and three plain tabs, having chords 10, 20, and 30 percent of the flap chord, were made. The purpose of these tests was to continue an investigation to supply structural and aerodynamic section data that may be applied to the design of horizontal and vertical tail surfaces. The results are presented as diagrams of resultant pressures and of resultant-pressure increments for the airfoil with the flap and the 20-percent-chord tab. Increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs are also given. At all unstalled flap and tab deflections, the experimental distributions agree well with those calculated by an analytical method. The agreement is poor, however, then the stalled or the unstalled condition of the flap or tab deflected alone was changed to an unstalled or stalled condition by the simultaneous deflection of both the flap and the tab.
Pressure-distribution tests of an NACA 0009 airfoil with an 80-percent-chord plain flap and three plain tabs, having chord of 10, 20, and 30 percent of the flap chord, were made. Section data suitable for application to the design of horizontal and vertical tail surfaces were obtained. Resultant-pressure diagrams for the airfoil with the flap and the 20-percent-chord tab are presented. Plots are also given of increments of normal-force and hinge-moment coefficients for the airfoil, the flap, and the three tabs. A comparison of some characteristic slopes for the 30-, the 50-, and the 80-percent-chord flaps, tested in the general investigation of plain flaps for control surfaces, is included. Section aerodynamic and load data have been made available for a wide range of flap and a tab chords to be used on an NACA 0009 airfoil or on other conventional sections.
A brief investigation of a target-type thrust reverser on a single-engine fighter model has been conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.05.At Mach numbers of 0.80, 0.92, and 1.05, a hydrogen peroxide turbojet-engine simulator was operated with the thrust reverser extended. The angle of attack was varied from 0 degrees to 5 degrees at these Mach numbers. The Reynolds number of the free stream, based on the mean aerodynamic chord, was about 5 x 10(6). It was estimated that reversed jet operations separated the model boundary-layer flow over the upper surface of the horizontal tail and upper part of the afterbody. This resulted in a positive pitch increment due to reversed jet operation. Jet-on operation also tended to stabilize the severe lateral oscillations which occurred with the reverser extended and the jet off. It appeared that these jet-off oscillations were the result of an alternating separation and reattachment of the flow on the rearmost portions of the fuselage afterbody.
Collection efficiency and ice accretion calculations were made for a commercial transport using the NASA Lewis LEWICE3D ice accretion code, the ICEGRID3D grid code and the CMARC panel code. All of the calculations were made on a Windows 95 based personal computer. The ice accretion calculations were made for the nose, wing, horizontal tail and vertical tail surfaces. Ice shapes typifying those of a 30 minute hold were generated. Collection efficiencies were also generated for the entire aircraft using the newly developed unstructured collection efficiency method. The calculations highlight the flexibility and cost effectiveness of the LEWICE3D, ICEGRID3D, CMARC combination.
Explore the source record for details and available documents.
An investigation has been conducted in the Langley stability tunnel on a vertical-tail model with a stub fuselage in combination with various horizontal tails to determine the effect of horizontal-tail span and vertical location of the horizontal tail relative to the vertical tail on the aerodynamic characteristics of an unswept tail assembly in sideslip. The results of the investigation indicated that the induced loading carried by the horizontal tail produced a rolling moment about the point of attachment to the vertical tail which was strongly influenced by horizontal-tail span and vertical locations. The greatest effect of horizontal-tail span on the rolling-moment derivative of the complete tail assembly was obtained for horizontal-tail locations near the top of the vertical tail. Span loadings which were reduced to the static-stability derivatives were calculated for each configuration tested by applying the well-known finite-step method used for wings to the intersecting surfaces of the vertical and horizontal tails. The finite-step method provides a simple and effective means of investigating the span loadings of intersecting surfaces.
Report presents the results of an investigation conducted in the Langley stability tunnel to determine the effects of horizontal tails of various sizes and at various tail lengths (when loaded on the fuselage center line) and also the effects of vertical location of the horizontal tail relative to the wing on the low-speed static longitudinal stability and on the steady-state rotary damping in pitch for a complete-model configuration. The wing and tail surfaces had the quarter-chord lines swept back 45 degrees and had aspect ratios of 4. The results of the investigation showed that, in agreement with analytical considerations, the contribution of the horizontal tail to static longitudinal stability was related directly to the tail size and length; whereas, its contribution to damping in pitch was related directly to tail size and the square of tail length.
Explore the source record for details and available documents.