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  • 1
    Electronic Resource
    Electronic Resource
    Solution of the axisymmetric inverse problem by higher-order line doublets (1994)
    Chichester : Wiley-Blackwell
    International Journal for Numerical Methods in Fluids 18 (1994), S. 415-432 
    Details
    ISSN: 0271-2091
    Keywords: Aerodynamics ; Aerodynamic design ; Inverse problems ; Body shaping ; Engineering ; Engineering General
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Mechanical Engineering, Materials Science, Production Engineering, Mining and Metallurgy, Traffic Engineering, Precision Mechanics
    Notes: The axial singularity inverse method for designing bodies of revolution has been improved by using higher-order doublet elements. The performance of the method for various element orders and other solution parameters is presented in detail. The results indicate that the method is generally more robust, less sensitive to insets and has a better-conditioned coefficient matrix compared with the source method of the same order. The condition number of the matrix is shown to increase with the thickness of the body, the order of the method, the number of elements and the degree of stretching of the node distribution. In general, good performance is attained for most bodies even with ƒr as low as 2 by using 10-12 second-order doublet elements with insets greater than 0.02L from rounded ends. Increasing the insets to 0.06L appears to improve the accuracy of the method for most bodies but slows its convergence.
    Additional Material: 9 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/fld.1650180406
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    Paper (German National Licenses)
    Fulltext
  • 2
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    Windsurfer Fin (1994)
    In:  CASI
    Details
    Publication Date: 2018-06-09
    Description: A NASA report detailing a wind tunnel investigation of a variable camber and twist could effectively reduce drag, thus improving performance. The resulting VooDoo fin is made of composite materials, has a rigid internal spar and a flexible polymer exterior coating. It is computer-designed and exceptionally durable.
    Keywords: Aerodynamics
    Type: Spinoff 1994; 79; NASA-NP-214
    Format: application/pdf
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  • 3
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    Simplified Aerodynamic and Structural Modeling for Oblique All-Wing Aircraft. Phase 2: Structures (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Any aircraft preliminary design study requires a structural model of the proposed configuration. The model must be capable of estimating the structural weight of a given configuration, and of predicting the deflections which will result from foreseen flight and ground loads. The present work develops such a model for the proposed Oblique All Wing airplane. The model is based on preliminary structural work done by Jack Williams and Peter Rudolph at Mdng, and is encoded in a FORTRAN program. As a stand-alone application, the program can calculate the weight CG location, and several types of structural deflections; used in conjunction with an aerodynamics model, the program can be used for mission analysis or sizing studies.
    Keywords: Aerodynamics
    Type: NASA-CR-202164 , NAS 1.26:202164
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  • 4
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    A supercritical airfoil experiment (1994)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference connections can be made to the data sets.
    Keywords: Aerodynamics
    Type: OTN-035236 , OTN-BIBL-AGARD-AR-303-Vol-2
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  • 5
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    Progress in Grid Generation: From Chimera to DRAGON Grids (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Hybrid grids, composed of structured and unstructured grids, combines the best features of both. The chimera method is a major stepstone toward a hybrid grid from which the present approach is evolved. The chimera grid composes a set of overlapped structured grids which are independently generated and body-fitted, yielding a high quality grid readily accessible for efficient solution schemes. The chimera method has been shown to be efficient to generate a grid about complex geometries and has been demonstrated to deliver accurate aerodynamic prediction of complex flows. While its geometrical flexibility is attractive, interpolation of data in the overlapped regions - which in today's practice in 3D is done in a nonconservative fashion, is not. In the present paper we propose a hybrid grid scheme that maximizes the advantages of the chimera scheme and adapts the strengths of the unstructured grid while at the same time keeps its weaknesses minimal. Like the chimera method, we first divide up the physical domain by a set of structured body-fitted grids which are separately generated and overlaid throughout a complex configuration. To eliminate any pure data manipulation which does not necessarily follow governing equations, we use non-structured grids only to directly replace the region of the arbitrarily overlapped grids. This new adaptation to the chimera thinking is coined the DRAGON grid. The nonstructured grid region sandwiched between the structured grids is limited in size, resulting in only a small increase in memory and computational effort. The DRAGON method has three important advantages: (1) preserving strengths of the chimera grid; (2) eliminating difficulties sometimes encountered in the chimera scheme, such as the orphan points and bad quality of interpolation stencils; and (3) making grid communication in a fully conservative and consistent manner insofar as the governing equations are concerned. To demonstrate its use, the governing equations are discretized using the newly proposed flux scheme, AUSM+, which will be briefly described herein. Numerical tests on representative 2D inviscid flows are given for demonstration. Finally, extension to 3D is underway, only paced by the availability of the 3D unstructured grid generator.
    Keywords: Aerodynamics
    Type: NASA-TM-106709 , NAS 1.15:106709 , ICOMP-94-19 , E-9071
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  • 6
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    Validation of the RPLUS3D Code for Supersonic Inlet Applications Involving Three-Dimensional Shock Wave-Boundary Layer Interactions (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A three-dimensional computational fluid dynamics code, RPLUS3D, which was developed for the reactive propulsive flows of ramjets and scramjets, was validated for glancing shock wave-boundary layer interactions. Both laminar and turbulent flows were studied. A supersonic flow over a wedge mounted on a flat plate was numerically simulated. For the laminar case, the static pressure distribution, velocity vectors, and particle traces on the flat plate were obtained. For turbulent flow, both the Baldwin-Lomax and Chien two-equation turbulent models were used. The static pressure distributions, pitot pressure, and yaw angle profiles were computed. In addition, the velocity vectors and particle traces on the flat plate were also obtained from the computed solution. Overall, the computed results for both laminar and turbulent cases compared very well with the experimentally obtained data.
    Keywords: Aerodynamics
    Type: NASA-TM-106579 , E-8839 , NAS 1.15:106579
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  • 7
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    Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.
    Keywords: Aerodynamics
    Type: NASA-TM-4542 , L-17272 , NAS 1.15:4542
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  • 8
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    The NASA Langley 8-foot Transonic Pressure Tunnel calibration (1994)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.
    Keywords: Aerodynamics
    Type: NASA-TP-3437 , L-17322 , NAS 1.60:3437
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  • 9
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    Navier-Stokes Calculations For a Venus Composition Probe With Ablation And Wake Radiation (1994)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: A numerical investigation is carried out to determine the magnitude of wake radiation for a proposed Venus composition probe. One of the scientific goals of the mission is to determine the atmospheric composition of Venus by examining the intensity of scattered sunlight through the wake of the vehicle during planetary entry. In the wake of the vehicle, excited particles generated in the bow shock and boundary layers absorb and emit radiation. Thus, the purpose of this study is to determine if the radiation sensor will be able to sense the incoming solar radiative flux relative to the radiative flux generated in the wake. During portions of the entry trajectory the incident surface heat flux will be high enough to produce significant ablation. Ablation products such as CN are known to be strong radiators. Also, the ablation will be driven by strong radiation emanating from the bow shock. Thus, radiation and ablation will be coupled into the Navier-Stokes flow solutions.
    Keywords: Aerodynamics
    Type: AIAA 29th Thermophysics Conference; Jun 19, 1995 - Jun 22, 1995; San Diego, CA; United States
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  • 10
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    Quiet Supersonic Wind Tunnel Development (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: The ability to control the extent of laminar flow on swept wings at supersonic speeds may be a critical element in developing the enabling technology for a High Speed Civil Transport (HSCT). Laminar boundary layers are less resistive to forward flight than their turbulent counterparts, thus the farther downstream that transition from laminar to turbulent flow in the wing boundary layer is extended can be of significant economic impact. Due to the complex processes involved experimental studies of boundary layer stability and transition are needed, and these are performed in "quiet" wind tunnels capable of simulating the low-disturbance environment of free flight. At Ames, a wind tunnel has been built to operate at flow conditions which match those of the HSCT laminar flow flight demonstration 'aircraft, the F-16XL, i.e. at a Mach number of 1.6 and a Reynolds number range of 1 to 3 million per foot. This will allow detailed studies of the attachment line and crossflow on the leading edge area of the highly swept wing. Also, use of suction as a means of control of transition due to crossflow and attachment line instabilities can be studied. Topics covered include: test operating conditions required; design requirements to efficiently make use of the existing infrastructure; development of an injector drive system using a small pilot facility; plenum chamber design; use of computational tools for tunnel and model design; and early operational results.
    Keywords: Aerodynamics
    Type: Aerospace Ground Test Facilities and Flight Testing XXIX Short Course; Apr 25, 1994 - May 05, 1994; Tullahoma, TN; United States
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  • 11
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    Wind Tunnel Testing for the Stratospheric Observatory for Infrared Astronomy (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: NASA Ames Research Center is pursuing the development of SOFIA, the Stratospheric Observatory For Infrared Astronomy. SOFIA will consist of a 2.5 meter telescope mounted aft of the wing of a Boeing 747 aircraft. Since a large portion of the infrared spectrum is not visible at ground level due to absorption by water vapor in the atmosphere below 40,000 feet, it is highly desirable to make observations above this altitude. SOFIA will provide the opportunity for astronomers to conduct high-altitude research for extended periods of time. Current study is focused on wind tunnel testing for the open cavity. If not controlled, air would create resonance and damage the telescope. For this reason, SOFIA will design a boundary layer control device to achieve laminar flow over the cavity. This also provides a clearer flow for seeing, thus improving resolution on infrared sources. Other effects being tested in the wind tunnel are aerodynamic torque loads on the telescope, and flutter loads on the tail.
    Keywords: Aerodynamics
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  • 12
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    Separated and Recovering Turbulent Boundary Layer Flow Behind a Backward Facing Step For Different Reynolds Numbers (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Experimental results for a two-dimensional separated turbulent boundary layer behind a backward facing step for five different Reynolds numbers are reported. Results are presented in the form of tables, graphs and a floppy disk for an easy access of the data. Reynolds number based on the step height was varied by changing the reference velocity upstream of the step, U(sub o), and the step height, h. Hot-wire measurement techniques were used to measure three Reynolds stresses and four triple-velocity correlations. In addition, surface pressure and skin friction coefficients were measured. All hot-wire measurements were acquired in a measuring domain which excluded recirculating flow region due to the directional insensitivity of hot-wires. The downstream extent of the domain from the step was 51 h for the largest and I 14h for the smallest step height. This significant downstream length permitted extensive study of the flow recovery. Prediction of perturbed flows and their recovery is particularly attractive for popular turbulence models since variations of turbulence length and time scales and flow interactions in different regions are generally inadequately predicted. The data indicate that the flow in the free shear layer region behaves like the plane mixing layer up to about 2/3 of the mean reattachment length when the flow interaction with the wall commences the flow recovery to that of an ordinary turbulent boundary layer structure. These changes of the flow do not occur abruptly with the change of boundary conditions. A reattachment region represents a transitional region where the flow undergoes the most dramatic adjustments to the new boundary conditions. Large eddies, created in the upstream free-shear layer region, are being torn, recirculated, reentrained back into the main stream interacting with the incoming flow structure. It is foreseeable that it is quite difficult to describe the physics of this region in a rational and quantitative manner other than statistical. Downstream of the reattachment point the flow recovers at different rates near the wall, in the newly developing internal boundary layer, and in the outer part of the flow. It appears that Reynolds stresses do not fully recover up to the longest recovery length of 114 h.
    Keywords: Aerodynamics
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  • 13
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    Analysis of F/A-18 Tail Buffet Data Acquired in the 80- by 120-Foot Wind Tunnel (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Tail buffet studies were conducted on a full-scale, production, F/A-18 fighter aircraft in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California. Tail buffet data were acquired over an angle-of-attack range of +20 deg to +40 deg, a side-slip range of -16 deg to + 16 deg, and at wind speeds up to 100 knots. The maximum speed corresponds to a Reynolds number of l2.3 x l0(exp 6) based on mean aerodynamic chord and a Mach number of 0. 15. The port, vertical tail fin was instrumented with ninety-six surface-pressure transducers, arranged in six by eight arrays, on each side of the fin. ne aircraft was also equipped with a removable Leading-Edge Extension (LEX) fence whose purpose is to reduce tail-buffet loads. Current analysis methods for the unsteady aerodynamic pressures and loads are described. Only results for the zero side-slip condition are to be presented, both with and without the LEX fence. Results of the time-averaged, power-spectral analysis are presented for the tail fin bending moments which are derived from the integrated pressure field. Local wave velocities on the tail surfaces are calculated from pressure correlations. It was found that the LEX fence significantly reduces the magnitude of the root-mean-square pressures and bending moments. Scaling and repeatability issues are addressed by comparing the present full scale results for pressures at the 60%-span and 45%-chord location with previous full-scale F/A-18 tail-buffet test in the 80- by 120- Foot Wind Tunnel, and with several small-scale tests. The comparisons show that the tail buffet frequency scales very well with tail chord and free-stream velocity, and that there is good agreement with the previous full-scale test. Root-mean-square pressures and power spectra do not scale as well as the frequency results. Addition of a LEX fence caused tail-buffet loads to be reduced at all model scales.
    Keywords: Aerodynamics
    Type: SAE Aerospace Atlantic Conference; Apr 18, 1994 - Apr 22, 1994; Dayton, OH; United States
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  • 14
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    Investigation of a Laminar Flow Leading Edge (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: The recent resurgence of interest in utilizing laminar flow on aircraft surfaces for reduction in skin friction drag has generated a considerable amount of research in natural laminar flow (NLF) and hybrid laminar flow control (HLFC) on transonic aircraft wings. This research has focused primarily on airfoil design and understanding transition behavior with little concern for the surface imperfections and manufacturing variations inherent to most production aircraft. In order for laminar flow to find wide-spread use on production aircraft, techniques for constructing the wings must be found such that the large surface imperfections present in the leading edge region of current aircraft do not occur. Toward this end, a modification to existing leading edge construction techniques was devised such that the resulting surface did not contain large gaps and steps as are common on current production aircraft of this class. A lowspeed experiment was first conducted on a simulation of the surface that would result from this construction technique. Preston tube measurements of the boundary layer downstream of the simulated joint and flow visualization using sublimation chemicals validated the literature on the effects of steps on a laminar boundary layer. These results also indicated that the construction technique was indeed compatible with laminar flow. In order to fully validate the compatibility of this construction technique with laminar flow, thus proving that it is possible to build wings that are smooth enough to be used on business jets and light transports in a manner compatible with laminar flow, a flight experiment is being conducted. In this experiment Mach number and Reynolds number will be matched in a real flight environment. The experiment is being conducted using the NASA Dryden F-104 Flight Test Fixture (FTF). The FTF is a low aspect ratio ventral fin mounted beneath an F-104G research aircraft. A new nose shape was designed and constructed for this experiment. This nose shape provides an accelerating pressure gradient in the leading edge region. By flying the aircraft at appropriate Mach numbers and altitudes, this nose shape simulates the leading edge region of a laminar flow wing for a business jet or light transport. Manufactured into the nose shape is a spanwise slot located approximately four inches downstream of the leading edge. The slot, which is an inch wide and one-eighth of an inch deep allows the simulation of surface imperfections, such as gaps and steps at skin joints, which will occur on aircraft using this new construction technique. By placing strips of aluminum of various sizes and shapes in the slot, the effect on the boundary layer of different sizes and shapes of steps and gaps will be examined. It is planned to use five different configurations, differing primarily in the size and number of gaps. Downstream of the slot, the state of the boundary layer is determined using hot film gages and Stanton gages. Agreement between these two very different techniques of measuring boundary layer properties is considered important to being able to state with confidence the effects on the boundary layer of the simulated manufacturing imperfections. To date, the aircraft has not flown. First flights of the aircraft are on schedule to begin October 4, 1993. Low-speed, preliminary experiments at matching Reynolds numbers have been completed.
    Keywords: Aerodynamics
    Type: AIAA 6th Biennial Flight Test Conference; Jun 20, 1994 - Jun 23, 1994; Colorado Springs, CO; United States
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  • 15
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    Adaptive Wing Camber Optimization: A Periodic Perturbation Approach (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Available redundancy among aircraft control surfaces allows for effective wing camber modifications. As shown in the past, this fact can be used to improve aircraft performance. To date, however, algorithm developments for in-flight camber optimization have been limited. This paper presents a perturbational approach for cruise optimization through in-flight camber adaptation. The method uses, as a performance index, an indirect measurement of the instantaneous net thrust. As such, the actual performance improvement comes from the integrated effects of airframe and engine. The algorithm, whose design and robustness properties are discussed, is demonstrated on the NASA Dryden B-720 flight simulator.
    Keywords: Aerodynamics
    Type: H-1998 , Automatic Control in Aerospace; 35-40|Aerospace Control; Sep 12, 1994 - Sep 16, 1994; Palo Alto, CA; United States
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  • 16
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    Unsteady Aerodynamic Behavior Measured in Hovering Flight (1994)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: A highly-instrumented UH-60A aircraft was tested at NASA-Ames Research Center from August 1993 to February 1994 obtaining an extensive data base for level flight, maneuvers, acoustics (both with respect to ground microphone arrays and inflight microphones), and flight dynamics. A majority of the data obtained are now in an electronic data base, however, only a small fraction of the data have been examined. The proposed paper will examine the issue of hovering steadiness in more detail. In particular, a single set of data obtained during ground acoustic testing may provide considerable insight as the wind speeds were measured at a hover height of 250 feet and the aircraft was positioned in 15 deg. steps in heading from 0 to 180 deg. Also, hover housekeeping data were obtained for many of the 31 flights and these will also allow a characterization of the unsteadiness. The variation in section lift will be examined in terms of the induced flow angle variation and this will be related to possible physical explanations.
    Keywords: Aerodynamics
    Type: AHS 51st Annual Forum and Technology Display; May 09, 1995 - May 11, 1995; Fort Worth, TX; United States
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  • 17
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    An Assessment of the Effect of Compressibility on Dynamic Stall (1994)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Compressibility plays a significant role in the development of separation on airfoils experiencing unsteady motion, even at moderately compressible free-stream flow velocities. This effect can result in completely changed stall characteristics compared to those observed at incompressible speed, and can dramatically affect techniques used to control separation. There has been a significant effort in recent years directed toward better understanding; of this process, and its impact on possible techniques for control of separation in this complex environment. A review of existing research in this area will be presented, with emphasis on the physical mechanisms that play such an important role in the development of separation on airfoils. The increasing impact of compressibility on the stall process will be discussed as a function of free-stream Mach number, and an analysis of the changing flow physics will be presented. Examples of the effect of compressibility on dynamic stall will be selected from both recent and historical efforts by members of the aerospace community, as well as from the ongoing research program of the present authors. This will include a presentation of a sample of high speed filming of compressible dynamic stall which has recently been created using real-time interferometry.
    Keywords: Aerodynamics
    Type: 33rd AIAA Aerospace Sciences Meeting; Jan 09, 1995 - Jan 12, 1995; Reno, NV; United States
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  • 18
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    Hypersonic Research Program at NASA Ames Research Center (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: This paper will describe the Airbreathing Hypersonic Research Program at NASA Ames Research Center. A main theme will be the "From Computation Through Flight" research effort. General research areas covered will include systems analysis, aerodynamics and aerothermodynamics, propulsion, materials, and flight research. Illustrative results from each discipline will be presented. The synergism between computational and experimental research will be demonstrated by examples. All examples given will have been published in the open literature.
    Keywords: Aerodynamics
    Type: AIAA Atmospheric Flight Mechanics Conference; Aug 01, 1994 - Aug 03, 1994; Scottsdale, AZ; United States
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  • 19
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    Aerodynamics Characteristics of Multi-Element Airfoils at -90 Degrees Incidence (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: A developed method has been applied to calculate accurately the viscous flow about airfoils normal to the free-stream flow. This method has special application to the analysis of tilt rotor aircraft in the evaluation of download. In particular, the flow about an XV-15 airfoil with and without deflected leading and trailing edge flaps at -90 degrees incidence is evaluated. The multi-element aspect of the method provides for the evaluation of slotted flap configurations which may lead to decreased drag. The method solves for turbulent flow at flight Reynolds numbers. The flow about the XV-15 airfoil with and without flap deflections has been calculated and compared with experimental data at a Reynolds number of one million. The comparison between the calculated and measured pressure distributions are very good, thereby, verifying the method. The aerodynamic evaluation of multielement airfoils will be conducted to determine airfoil/flap configurations for reduced airfoil drag. Comparisons between the calculated lift, drag and pitching moment on the airfoil and the airfoil surface pressure will also be presented.
    Keywords: Aerodynamics
    Type: AIAA Aerospace Sciences; Jan 09, 1995 - Jan 12, 1995; Reno, NV; United States
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  • 20
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    Navier-Stokes Flowfield Simulation of Boeing 747-200 as Platform for SOFIA (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Steady and unsteady viscous, three-dimensional flowfields are calculated using a thin layer approximation of Navier-Stokes equations in conjunction with Chimera overset grids. The finite-difference numerical scheme uses structured grids and a pentadiagonal flow solver called "OVERFLOW". The configuration of Boeing 747-200 has been chosen as one of configurations to be used as a platform for the SOFIA (Stratospheric Observatory For Infrared Astronomy). Initially, the steady flowfield of the full aircraft is calculated for the clean configuration (without a cavity to house telescope). This solution is then used to start the unsteady flowfield of a configuration containing cavity housing the observation telescope and its peripheral units. Analysis of unsteady flowfield in the cavity and its influence on the tail empennage, as well as the noise due to turbulence and optical quality of the flow are the main focus of this study. For the configuration considered here, the telescope housing cavity is located slightly downstream of the portwing. The entire flow-field is carefully constructed using 45 overset grids and consists of nearly 4 million grid points. All the computations axe done at one freestream flow condition of M(sub infinity) = 0.85, alpha = 2.5deg, and a Reynolds of Re = 1.85x10deg
    Keywords: Aerodynamics
    Type: AIAA Aerspace Sciences; Jan 02, 1995 - Jan 12, 1995; Reno, NV; United States
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  • 21
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    Wind Tunnel Visualization of the Flow Over a Full-Scale F/A-18 Aircraft (1994)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: The proposed paper presents flow visualization performed during experiments conducted on a full-scale F/A-18 aircraft in the 80- by 120-Foot Wind-Tunnel at NASA Ames Research Center. This investigation used both surface and off-surface flow visualization techniques to examine the flow field on the forebody, canopy, leading edge extensions (LEXs), and wings. The various techniques used to visualize the flow field were fluorescent tufts, flow cones treated with reflective material, smoke in combination with a laser light sheet, and a video imaging system. The flow visualization experiments were conducted over an angle of attack range from 20deg to 45deg and over a sideslip range from -10deg to 10deg. The results show regions of attached and separated flow on the forebody, canopy, and wings. Additionally, the vortical flow is clearly visible over the leading-edge extensions, canopy, and wings.
    Keywords: Aerodynamics
    Type: SAE Aerospace Atlantic Conference; Apr 18, 1994 - Apr 22, 1994; Dayton, OH; United States
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  • 22
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    Comment on 'Guidance for an Aerocapture Maneuver (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: It is stated that the aerodynamic forces on the vehicle being aerocaptured are controlled by "altering the angle of attack" and thereby controlling the lift coefficient. Furthermore, the resulting variation of drag coefficient with angle of attack was ignored. The purpose of this Comment is to point out that an aerodynamic control method that is much more effective than the pitch modulation has been studied and utilized during entries for many years. During aerocapture, it is desirable to have a large range of lift coefficients available, while keeping the vehicle's ballistic coefficients constant. This is accomplished by modulating the vehicle's bank angle, i.e., by rolling the vehicle about its velocity vector. By this method, the angle of attack can be held constant (at the trim angle, if desired), and the C(sub D) and the ballistic coefficient remain constant. Furthermore, the vertical component of the normal force vector (essentially the lift) can be varied over its entire range, from maximum positive to maximum negative values. Reaction controls, rather than aerodynamic ones, are usually utilized to change the bank angle of the vehicle, thus requiring the use of fuel. However, the fuel expenditure that is required to change the bank angle is far less than the amount that would have to be used to continuously hold the vehicle at pitch angles that differ significantly from its trim angle of attack. Also, it has been shown that bank angle modulation to vary the lift can enlarge the entry corridor by increasing the entry angle for the undershoot boundary, where both the heating rate and deceleration reach a maximum. Finally, the crew's deceleration tolerance can be increased somewhat when the bank angle is varied, as opposed to the pitch angle. For bank modulation, the deceleration force vector can be kept at a constant angle with respect to the occupants whose tolerance to g loads is highest when the force is applied in a direction normal to the upper torso. The advantages of bank angle variation to modulate the lift vector were recognized long ago, and this method of control was used successfully on the Apollo command module during lunar return' and, more recently, for the Space Shuttle Orbiter.
    Keywords: Aerodynamics
    Type: Journal of Guidance, Control, and Dynamics; 17; 4; 878-878
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  • 23
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    Experimental and Computational Study of Sonic and Supersonic Jet Plumes (1994)
    In:  Other Sources
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    Publication Date: 2019-07-18
    Description: Study of sonic and supersonic jet plumes are relevant to understanding such phenomenon as jet-noise, plume signatures, and rocket base-heating and radiation. Jet plumes are simple to simulate and yet, have complex flow structures such as Mach disks, triple points, shear-layers, barrel shocks, shock-shear-layer interaction, etc. Experimental and computational simulation of sonic and supersonic jet plumes have been performed for under- and over-expanded, axisymmetric plume conditions. The computational simulation compare very well with the experimental observations of schlieren pictures. Experimental data such as temperature measurements with hot-wire probes are yet to be measured and will be compared with computed values. Extensive analysis of the computational simulations presents a clear picture of how the complex flow structure develops and the conditions under which self-similar flow structures evolve. From the computations, the plume structure can be further classified into many sub-groups. In the proposed paper, detail results from the experimental and computational simulations for single, axisymmetric, under- and over-expanded, sonic and supersonic plumes will be compared and the fluid dynamic aspects of flow structures will be discussed.
    Keywords: Aerodynamics
    Type: AIAA Atmospheric Flight Mechanics Conference; Aug 07, 1995 - Aug 09, 1995; Baltimore, MD; United States
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  • 24
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    The Effect of Two-Dimensional Forcing on Self-Similar Turbulent Plane Wakes (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Three direct numerical simulations of time-evolving turbulent plane wakes with velocity deficit Reynolds numbers of about 2,000 have been simulated using a spectral numerical method with up to 600 x 260 x 160 modes. The initial conditions for the simulations are generated from direct numerical simulations of a turbulent boundary layer (momentum thickness Reynolds number of 670), and varying amounts of additional two- dimensional, forcing. In order to preserve the self-similar flow evolution, the forcing is implemented by multiplying all the two-dimensional modes in the initial condition by a constant factor. In the "natural" case no additional forcing is used; in the "forced" and "heavily forced" cases this factor is 5 and 20, respectively. The wake spreading rate Is increased by factors of 1.7 and 7.1 for the two forced cases. The Reynolds stresses are also increased by a similar or even larger factor. These results indicate that the plane wake is much more sensitive to initial forcing than the plane mixing layer. As in the plane mixing layer, two-dimensional forcing promotes more organized large-scale vortical flow structures and these structures axe sometimes separated by "braid regions" containing streamwise "rib" vortices, unlike in the unforced wake.
    Keywords: Aerodynamics
    Type: Forty-Seventh Annual Meeting of the American Physical Society; Nov 20, 1994 - Nov 22, 1994; Atlanta, GA; United States
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  • 25
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    On the Two-Dimensionalization of Homogeneous Rotating Turbulence (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: Large-eddy simulation of the incompressible Navier-Stokes equations has been used to examine the long-time development of initially isotropic turbulence subjected to solid-body rotation. The simulations were carried out using a pseudo-spectral method with 128 x 128 x 512 collocation points in a computational domain that is four times larger along the rotation axis than in the other directions; subgrid-scale motions were parameterized using a spectral eddy viscosity model modified for system rotation. Simulation results show that the correlation length along the rotation am's of velocities orthogonal to the rotation vector exhibits rapid growth while the integral length-scale of velocities aligned with the rotation axis is relatively unaffected by rotation. Examination of the energy spectrum of two-dimensional, two-component motions indicates the presence of an inverse cascade of energy. System rotation also causes an alignment of vorticity along the rotation axis with relatively stronger cyclonic vorticity than anticyclonic. The onset of anisotropic effects are well characterized by Rossby numbers defined in terms of both macroscopic and microscopic quantities.
    Keywords: Aerodynamics
    Type: Forty-Seventh Annual Meeting of the American Physical Society; Nov 20, 1994 - Nov 22, 1994; Atlanta, GA; United States
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  • 26
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    Navier-Stokes Solutions of Rotor and Rotor-Body Flows (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: This paper will review the advances made recently in the Navier-Stokes CFD methods to simulate aerodynamics and aeroacoustics of helicopter rotors and rotor-body flows. Although a complete flowfield simulation of full helicopter is currently not feasible with these methods, impressive gains have been made in analyzing individual components of this complex problem in a very detailed manner. The use of the state-of-the-art numerical algorithms in solution methods, in conjunction with powerful supercomputers, like the Cray-2, have enabled noticeable progress to be made in modeling viscous-inviscid interactions, blade-vortex interactions, tip-vortex: simulation and wake effects, as well as high speed impulsive noise in hover and forward flight for isolated rotor blades. This paper will critically evaluate the presently available Euler and Navier-Stokes methods, both finite-difference and finite volume methods using structured and unstructured grids for helicopter applications for accuracy, suitability, and computational efficiency. The review will also include the recent progress made using overset grids to model rotor-body flows. All the material for this review will be drawn from the published material shown below.
    Keywords: Aerodynamics
    Type: International Colloquium on Vortical Flows in the Aeronautics; Oct 12, 1994 - Oct 14, 1994; Aachan; Germany
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  • 27
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    Wing-Body Aeroelasticity Using Finite-Difference Fluid/Finite-Element Structural Equations on Parallel Computers (1994)
    In:  Other Sources
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    Publication Date: 2019-07-17
    Description: In recent years significant advances have been made for parallel computers in both hardware and software. Now parallel computers have become viable tools in computational mechanics. Many application codes developed on conventional computers have been modified to benefit from parallel computers. Significant speedups in some areas have been achieved by parallel computations. For single-discipline use of both fluid dynamics and structural dynamics, computations have been made on wing-body configurations using parallel computers. However, only a limited amount of work has been completed in combining these two disciplines for multidisciplinary applications. The prime reason is the increased level of complication associated with a multidisciplinary approach. In this work, procedures to compute aeroelasticity on parallel computers using direct coupling of fluid and structural equations will be investigated for wing-body configurations. The parallel computer selected for computations is an Intel iPSC/860 computer which is a distributed-memory, multiple-instruction, multiple data (MIMD) computer with 128 processors. In this study, the computational efficiency issues of parallel integration of both fluid and structural equations will be investigated in detail. The fluid and structural domains will be modeled using finite-difference and finite-element approaches, respectively. Results from the parallel computer will be compared with those from the conventional computers using a single processor. This study will provide an efficient computational tool for the aeroelastic analysis of wing-body structures on MIMD type parallel computers.
    Keywords: Aerodynamics
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  • 28
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    Cluster Computation of Flight Reynolds Number Flows (1994)
    In:  Other Sources
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    Publication Date: 2019-08-13
    Description: The performance of a workstation cluster used for the solution of the Reynolds-averaged Navier-Stokes equations is compared with a conventional vector supercomputer architecture. The application simulation of the steady flowfield about a transonic transport was computed using an implicit diagonal scheme in an overset mesh framework. Static load balancing was used, while coarse grain decomposition was achieved by solution of a grid zone per processor. Price/performance ratios are estimated for several scenarios in which such clusters may be utilized.
    Keywords: Aerodynamics
    Type: OAI/NASA Symposium Application of Parallel and Distributed Computing; Apr 18, 1994 - Apr 19, 1994; Columbus, OH; United States
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  • 29
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    Off-Axis Drag of Dendrite Fragments at Low Reynolds Number (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The aim of the present investigation is to characterize the motion of dendrite fragments falling under the influence of gravity in a uniform liquid medium at low Reynolds number. In an earlier study, Zakhem, Weidman and de Groh (1992) reported on the settling speed of model equiaxed dendrite grains released along their axis of symmetry. In this follow-up study uniaxial model dendrite grains were released off-axis to observe and document their motion at different orientations. It was hypothesized that the dendrite models might rotate when released off-axis in which case an attempt would be made to document the ensuing unsteady motion. This latter event turned out to be in fact true: at the small but finite Reynolds numbers that existed, each uniaxial dendrite slowly rotated towards its equilibrium orientation while failing under the influence of gravity. In addition to completing the original goal, we have made use of a beads-on-a shell Stokes flow code to numerically determine the drag coefficient for capsules, i.e.. uniaxial dendrites without arms. The drag on horizontally and vertically falling capsules are reported and compared with measurements.
    Keywords: Aerodynamics
    Type: NASA/CR-94-207107 , NAS 1.26:207107
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  • 30
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    Investigation of supersonic jet plumes using an improved two-equation turbulence model (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Supersonic jet plumes were studied using a two-equation turbulence model employing corrections for compressible dissipation and pressure-dilatation. A space-marching procedure based on an upwind numerical scheme was used to solve the governing equations and turbulence transport equations. The computed results indicate that two-equation models employing corrections for compressible dissipation and pressure-dilatation yield improved agreement with the experimental data. In addition, the numerical study demonstrates that the computed results are sensitive to the effect of grid refinement and insensitive to the type of velocity profiles used at the inflow boundary for the cases considered in the present study.
    Keywords: Aerodynamics
    Type: NASA-TM-111555 , NAS 1.15:111555 , AIAA Paper 92-2604 , Applied Aerodynamics Conference; Jun 22, 1992 - Jun 24, 1992; Palo Alto, CA; United States
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  • 31
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    Viscous Shock-Layer Analysis of Two-Dimensional and Axisymmetric Flows (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a two-dimensional viscous shock-layer code for nonequilibrium gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. This study includes results for 6-deg half-angle bodies with nose radii ranging from 0.01 to 2.0 ft for both cylindrically blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. Generally, this ratio and the individual heating rates are smaller for cylindrically blunted wedges with small nose radii as compared to the spherically blunted cones (for the same nose radius). Therefore, a larger potential exists for heating reduction in cylindrically blunted as compared with the spherically blunted surfaces. However, the results presented at higher altitudes (where the slip effects become important) show that the spherically, blunted nose gives lower stagnation-point heating due to stronger merged shock-layer effects as compared with a cylindrically blunted nose.
    Keywords: Aerodynamics
    Type: NASA-TM-111564 , NAS 1.15:111564 , AIAA Paper 93-2751 , Thermophysics Conference; Jul 06, 1993 - Jul 09, 1993; Orlando, FL; United States
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  • 32
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    Navier-Stokes prediction of large-amplitude forced and free-to-roll delta-wing oscillations (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The three-dimensional Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate nonsteady vortical flow about a 65 degree sweep delta wing at 30 degrees angle of attack. Two large-amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental forces, moments, and roll-angle time histories. The overall agreement is good. Vortex breakdown is present in each case, which causes significant time lags in the vortex breakdown motions relative to the body motions. This behavior strongly influences the dynamic forces and moments.
    Keywords: Aerodynamics
    Type: NASA-TM-111611 , NAS 1.15:111611 , AIAA Paper 94-1884 , AIAA Applied Aerodynamics Conference; Jun 20, 1994 - Jun 22, 1994; Colorado Springs, CO; United States
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  • 33
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    Validation of Vortex-Lattice Method for loads on wings in lift-generated wakes (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A study is described that evaluates the accuracy of vortex-lattice methods when they are used to compute the loads induced on aircraft as they encounter lift-generated wakes. The evaluation is accomplished by use of measurements made in the 80- by 120-foot wind tunnel of the lift, rolling-moment, and downwash in the wake of three configurations of a model of a subsonic transport aircraft. The downwash measurements are used as input for a vortex-lattice code in order to compute the lift and rolling moment induced on wings that have a span of 0.186, 0.510, or 1.022 times the span of the wake-generating model. Comparison of the computed results with the measured lift and rolling moment distributions are used to determine the accuracy of the vortex-lattice code. It was found that the vortex-lattice method is very reliable as long as the span of the encountering of following wing is less than about 0.2 of the generator span. As the span of the following wing increases above 0.2, the vortex-lattice method continues to correctly predict the trends and nature of the induced loads, but it overpredicts the magnitude of the loads by increasing amounts. The increase in deviation of the computed from the measured loads with size of the following wing is attributed to the increase in distortion of the structure of the vortex wake as it approaches and passes the larger following wings.
    Keywords: Aerodynamics
    Type: NASA-TM-111610 , NAS 1.15:111610 , AIAA-94-1839 , AIAA Applied Aerodynamics Conference; Jun 20, 1994 - Jun 22, 1994; Colorado Springs, CO; United States
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  • 34
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    Navier-Stokes Calculations for a Highly-Twisted Rotor Near Stall (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The viscous flow field near the surface of a hovering rotor blade was studied for blade twist distributions typical of a till rotor blade and a conventional helicopter rotor blade. Three blade geometries were studied, including a tilt rotor blade twist distribution (baseline), conventional helicopter rotor blade twist distribution, and the baseline twist distribution with 2 deg of precone. The results give insight into the delayed stall phenomenon often observed for highly twisted rotors. Calculations were performed for a high thrust condition near stall using the thin-layer Navier-Stokes CFD code TURNS. Effects of built-in twist on section force coefficients, skin friction, velocities, surface pressures, and boundary layer shape factor are discussed. Although the rotor thrust coefficient was nominally the same for the cases using the two twist distributions, large differences were found in the section in-plane and normal force coefficients. These preliminary results imply that the blade outboard region, rather than the inboard region, provides the majority of the performance advantage of the baseline case over the low twist case. Skin friction, velocities near the blade, and surface pressures for the two twist distributions reveal significant differences in the blade outboard region.
    Keywords: Aerodynamics
    Type: NASA-TM-111741 , NAS 1.15:111741 , Aeromechanics Specialists; Jan 19, 1994 - Jan 21, 1994; San Fransisco, CA; United States
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  • 35
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    A hybrid Navier-Stokes/full-potential method for the prediction of iced wing aerodynamics (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A hybrid method for computing compressible viscous flows is presented. This method divides the computational domain into two zones. In the outer zone, the unsteady full-potential equation (FPE) is solved. In the inner zone, the Navier-Stokes equations are solved. The two zones are tightly coupled so that steady and unsteady flows may be efficiently solved. The resulting CPU times are less than 50 percent of the required for a full-blown Navier-Stokes analysis. Sample applications of the method to an unswept iced wing at 4 deg and 8 deg angle of attack are presented. Surface pressures are in good agreement with the measurements obtained by Bragg et al. at the University of Illinois.
    Keywords: Aerodynamics
    Type: NASA-CR-201432 , NAS 1.26:201432 , AIAA Paper 94-0489 , AIAA Aerospace Sciences Meeting and Exhibit; Jan 10, 1994 - Jan 13, 1994; Reno, NV; United States
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  • 36
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    Static aeroelastic analysis of wings using Euler/Navier-Stokes equations coupled with improved wing-box finite element structures (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Strong interactions between flow about an aircraft wing and the wing structure can result in aeroelastic phenomena which significantly impact aircraft performance. Time-accurate methods for solving the unsteady Navier-Stokes equations have matured to the point where reliable results can be obtained with reasonable computational costs for complex non-linear flows with shock waves, vortices and separations. The ability to combine such a flow solver with a general finite element structural model is key to an aeroelastic analysis in these flows. Earlier work involved time-accurate integration of modal structural models based on plate elements. A finite element model was developed to handle three-dimensional wing boxes, and incorporated into the flow solver without the need for modal analysis. Static condensation is performed on the structural model to reduce the structural degrees of freedom for the aeroelastic analysis. Direct incorporation of the finite element wing-box structural model with the flow solver requires finding adequate methods for transferring aerodynamic pressures to the structural grid and returning deflections to the aerodynamic grid. Several schemes were explored for handling the grid-to-grid transfer of information. The complex, built-up nature of the wing-box complicated this transfer. Aeroelastic calculations for a sample wing in transonic flow comparing various simple transfer schemes are presented and discussed.
    Keywords: Aerodynamics
    Type: NASA-CR-201433 , NAS 1.26:201433 , AIAA Paper 94-1587 , AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference; Apr 18, 1994 - Apr 20, 1994; Hilton Head, SC; United States
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  • 37
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    Rarefield-Flow Shuttle Aerodynamics Flight Model (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.
    Keywords: Aerodynamics
    Type: NASA-TM-111566 , NAS 1.15:111566 , AIAA Paper 93-3441 , Applied Aerodynamics Conference; Aug 09, 1993 - Aug 11, 1993; Monterey, CA; United States
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  • 38
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    Analysis of tangential slot blowing on F/A-18 isolated forebody (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and under-blowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind tunnel and sub-scale wind tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flow field about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.
    Keywords: Aerodynamics
    Type: NASA-TM-111696 , NAS 1.15:111696 , AIAA Paper 95-1831 , AIAA Applied Aerodynamics Conference; Jun 20, 1994 - Jun 23, 1994; Colorado Springs, CO; United States
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  • 39
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    Measurements of Supersonic Wing Tip Vortices (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.
    Keywords: Aerodynamics
    Type: NASA-CR-202591 , NAS 1.26:202591 , AIAA Paper 94-2576 , Aerospace Ground Testing; Jun 20, 1994 - Jun 23, 1994; Colorado Springs, CO; United States
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  • 40
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    Severe Turbulence and Maneuvering from Airline Flight Records (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Digital flight records from reported clear-air turbulence incidents are used to determine winds and turbulence, to determine maneuver g loads, and to analyze control problems. Many cases of severe turbulence are found downwind of mountains and thunderstorms where sharp, sudden jolts are associated with vortices in atmospheric waves. Other cases of severe turbulence are round in strong updrafts above thunderstorm buildups that may be undetected by onboard weather radar. An important finding is that there are large maneuvering loads in over half of the reported clear-air turbulence incidents. Maneuvering loads are determined through an analysis of the short-term variations in elevator deflection and aircraft pitch angle. For altitude control in mountain waves the results indicate that small pitch angle changes with proper timing are sufficient to counter variations in vertical wind. For airspeed control in strong mountain waves, however, there is neither the available thrust nor the quickness in engine response necessary to counter the large variations in winds.
    Keywords: Aerodynamics
    Type: NASA-TM-111780 , NAS 1.15:111780 , AIAA Paper 92-4341 , Journal of Aircraft; 31; 4; 753-760|Atmospheric Flight Mechanics; Aug 10, 1992 - Aug 12, 1992; Hilton Head, SC; United States
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  • 41
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    Forebody Tangential Slot Blowing on an Aircraft Geometry (1994)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The effect of forebody tangential slot blowing on the flowfield about an F/A-18 aircraft is investigated numerically using solutions of the Navier-Stokes equations. Computed solutions are obtained for an aircraft geometry which includes the fuselage, a wing with deflected leading-edge flap, empennage, and a faired-over engine inlet. The computational slot geometry corresponds to that used in full-scale wind-tunnel tests. Solutions are computed using flight test conditions and jet mass flow ratios equivalent to wind-tunnel test conditions. The effect of slot location is analyzed by computing two nontime-accurate solutions with a 16-in. slot located 3 in. and 11 in. aft of the nose of the aircraft. These computations resolve the trends observed in the full-scale wind-tunnel test data. The flow aft of the leading-edge extension vortex burst is unsteady. A time-accurate solution is obtained to investigate the flow characteristics aft of the vortex burst, including the effect of blowing on tail buffet.
    Keywords: Aerodynamics
    Type: NASA-TM-111779 , NAS 1.15:111779 , AIAA Paper 93-2962 , Journal of Aircraft; 31; 4; 922-928|Fluid Dynamics, Plasmadynamics, and Lasers Conference; Jul 06, 1993 - Jul 09, 1993; Orlando, Fl; United States
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    An Exploratory Study of a Subspace Iteration Method as an Alternative to the QR Method for Floquet Eigenanalysis (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Floquet eigenanalysis requires a few dominant eigenvalues of the Floquet transition matrix (FTM). Although the QR method is used almost exclusively, it is expensive for such partial eigenanalysis; the operation counts and, thereby, the approximate machine-time grow cubically with the matrix order. Accordingly, for Floquet eigenanalysis, the Arnold-Saad method, a subspace iteration method, is investigated as an alternative to the QR method. The two methods are compared for machine-time efficiency and the residual errors of the corresponding eigenpairs. The Arnolds-Saad method takes much less machine-time than the QR method with comparable computational reliability and offers promise fpr large-scale Floquet eigenanalysis.
    Keywords: Aerodynamics
    Type: NASA-CR-203147 , NAS 1.26:203147 , Mathl. Comput. Modelling (ISSN 0895-7177); 19; 4-Mar; 69-73
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    Application of a Third Order Upwind Scheme to Viscous Flow over Clean and Iced Wings (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A 3-D compressible Navier-Stokes solver has been developed and applied to 3-D viscous flow over clean and iced wings. This method uses a third order accurate finite volume scheme with flux difference splitting to model the inviscid fluxes, and second order accurate symmetric differences to model the viscous terms. The effects of turbulence are modeled using a Kappa-epsilon model. In the vicinity of the sold walls the kappa and epsilon values are modeled using Gorski's algebraic model. Sampling results are presented for surface pressure distributions, for untapered swept clean and iced wings made of NACA 0012 airfoil sections. The leading edge of these sections is modified using a simulated ice shape. Comparisons with experimental data are given.
    Keywords: Aerodynamics
    Type: NASA-CR-202616 , NAS 1.26:202616 , AIAA Paper 94-0485 , Aerospace Sciences; Jan 10, 1994 - Jan 13, 1994; Reno, NV; United States
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    An Experimental Simulation of Flap Flow on Multielement Airfoils at High Reynolds Number (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The design of the high lift system has a major impact on the performance of an aircraft yet our understanding of the physics of this flow is still weak. Flow features include interactions between the wakes shed from the upstream elements and the pressure gradients and boundary layers of the downstream elements. Interaction of the turbulent wake of the main element and the flap can cause (1) separation of the flap boundary layer or (2) 'bursting' of the main airfoil wake. Although the first factor is at least partially understood, even the qualitative aspects of (2) remain to be determined. In order to study these phenomena at Reynolds numbers approaching those of flight, a thick high Reynolds number wake is created using a 24 foot flat plate in the long rectangular test section of a 4 ft. by 6 ft subsonic wind tunnel. The design and construction of this test section, plate, and accompanying flap is described. Results obtained in a quarter-scale model were used for design purposes and are also described. Construction of the full scale facility is complete and preliminary results are presented.
    Keywords: Aerodynamics
    Type: NASA-CR-203019 , NAS 1.26:203019 , AIAA Paper 94-2613 , Aerospace Ground Testing; Jun 20, 1994 - Jun 23, 1994; Colorado Springs, CO; United States
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    A Review of Evidence for High Life Coefficients on Propeller and Rotor Blades Under Static Thrust Conditions with Some New Experimental Results (1994)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Interest has increased recently in the thrust-producing capability of rotors at very high collective pitch angles. An early reference noted this behaviour in rotors and offered alternative models for section lift characteristics to explain it. The same phenomenon was coincidentally noted and used in a propeller code, resulting in very good correlation with static thrust data. The proposed paper will present experimental data demonstrating the pronounced persistence of thrust for propellers at increasing collective pitch angles. Comparisons with blade element/momentum theory will be made. These results are expected to point to the need to define (ultimately to explain) aerodynamic lift and drag behaviour in a rotating environment. Experimental measurements made by the U.S. Army Aeroflightdynamics Directorate at the Ames Research Center have shown that locally measured normal force coefficients along the span of a highly twisted rotor blade continue to increase at high values of collective pitch. In some cases these coefficients exceed expected values for the same type of airfoil tested under two dimensional conditions. To date no one to the authors' knowledge has defined the variation of C(n) with pitch for very high angles (to 45 deg) in a rotating environment and for a blade of reasonably high aspect ratio; however, total propeller thrust measurements support the idea that stalling does not occur in the same way as on a wing. This paper will present experimental data in the form of surface pressure distributions as well as flow visualization (microtufts) to explore the aerodynamic behavior of the rotating airfoil at high values of blade incidence. This paper also reviews experimental evidence and infers some high lift coefficient behavior from it. Comparisons between predicted thrust, utilizing modified airfoil characteristics and a blade element model, and measured thrust for both rotors and propellers that cover the extremes of collective pitch are shown and discussed.
    Keywords: Aerodynamics
    Type: AGARD Aerodynamic Conference; Oct 01, 1994; Berlin; Germany
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    Sensitivity Analysis of the Static Aeroelastic Response of a Wing (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A technique to obtain the sensitivity of the static aeroelastic response of a three dimensional wing model is designed and implemented. The formulation is quite general and accepts any aerodynamic and structural analysis capability. A program to combine the discipline level, or local, sensitivities into global sensitivity derivatives is developed. A variety of representations of the wing pressure field are developed and tested to determine the most accurate and efficient scheme for representing the field outside of the aerodynamic code. Chebyshev polynomials are used to globally fit the pressure field. This approach had some difficulties in representing local variations in the field, so a variety of local interpolation polynomial pressure representations are also implemented. These panel based representations use a constant pressure value, a bilinearly interpolated value. or a biquadraticallv interpolated value. The interpolation polynomial approaches do an excellent job of reducing the numerical problems of the global approach for comparable computational effort. Regardless of the pressure representation used. sensitivity and response results with excellent accuracy have been produced for large integrated quantities such as wing tip deflection and trim angle of attack. The sensitivities of such things as individual generalized displacements have been found with fair accuracy. In general, accuracy is found to be proportional to the relative size of the derivatives to the quantity itself.
    Keywords: Aerodynamics
    Type: NASA-CR-200793 , NAS 1.26:200793
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    Robust Crossfeed Design for Hovering Rotorcraft (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Control law design for rotorcraft fly-by-wire systems normally attempts to decouple angular responses using fixed-gain crossfeeds. This approach can lead to poor decoupling over the frequency range of pilot inputs and increase the load on the feedback loops. In order to improve the decoupling performance, dynamic crossfeeds may be adopted. Moreover, because of the large changes that occur in rotorcraft dynamics due to small changes about the nominal design condition, especially for near-hovering flight, the crossfeed design must be 'robust'. A new low-order matching method is presented here to design robust crossfeed compensators for multi-input, multi-output (MIMO) systems. The technique identifies degrees-of-freedom that can be decoupled using crossfeeds, given an anticipated set of parameter variations for the range of flight conditions of concern. Cross-coupling is then reduced for degrees-of-freedom that can use crossfeed compensation by minimizing off-axis response magnitude average and variance. Results are presented for the analysis of pitch, roll, yaw and heave coupling of the UH-60 Black Hawk helicopter in near-hovering flight. Robust crossfeeds are designed that show significant improvement in decoupling performance and robustness over nominal, single design point, compensators. The design method and results are presented in an easily used graphical format that lends significant physical insight to the design procedure. This plant pre-compensation technique is an appropriate preliminary step to the design of robust feedback control laws for rotorcraft.
    Keywords: Aerodynamics
    Type: NASA-CR-202403 , NAS 1.26: 202403
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    Hover Acoustic Characteristics of the XV-15 with Advanced Technology Blades (1993)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experiment has been performed to investigate the far-field hover acoustic characteristics of the XV-15 aircraft with advanced technology blades (ATB). An extensive, high-quality, far-field acoustics data base was obtained for a rotor tip speed range of 645-771 ft/s. A 12-microphone, 500-ft radius semicircular array combined with two aircraft headings provided acoustic data over the full 360-deg azimuth about the aircraft with a resolution of 15 deg. Altitude variations provided data from near in-plane to 45 deg below the rotor tip path plane. Acoustic directivity characteristics in the lower hemisphere are explored through pressure time histories, narrow-band spectra, and contour plots. Directivity patterns were found to vary greatly with azimuth angle, especially in the forward quadrants. Sharp positive pressure pulses typical of blade-vortex interactions were found to propagate aft of the aircraft and were most intense at 45 deg below the rotor plane. Modest overall sound pressure levels were measured near in-plane indicating that thickness noise is not a major problem for this aircraft when operating in the hover mode with ATB. Rotor tip speed reductions reduced the average overall sound pressure level (dB (0.0002 dyne/cm(exp 2)) by nearly 8 dB in-plane, and 12.6 deg below the rotor plane.
    Keywords: Aerodynamics
    Type: NASA-TM-111578 , NAS 1.15:111578
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    A Numerical Simulation Study to Develop an Acceptable Wake Encounter Boundary for a B737-100 Airplane (1993)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: The National Aeronautics and Space Administration (NASA) is conducting research with the goal of enabling safe improvements in the capacity of the nation's air transportation system. The wake-vortex upset hazard is an important factor in establishing the minimum safe spacing between aircraft during landing and take-off operations, thus impacting airport capacity. A batch simulation study was conducted to assess the sensitivity of various safe landing criteria in the development of an acceptable wake encounter boundary. A baseline six-degree-of-freedom simulation of a B737-100 airplane was modified to include a wake model and the vortex-induced forces and moments. The guidance and control input for the airplane was provided by an auto-land system. The wake strength and encounter geometry were varied. A sensitivity study was also conducted to assess the effects of encounter modeling methods and accuracy.
    Keywords: Aerodynamics
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    Large-Eddy Simulation of Boundary Layer Transition on Swept Wings (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The large-eddy simulation of the spatial evolution of a stationary crossflow vortex packet in a three-dimensional boundary layer was performed. Although a coarse grid was used (compared to that required by a direct numerical simulation) the essential features of the disturbance evolution, such as the spanwise disturbance spreading and the vortex rollover, were captured accurately. The eddy viscosity became significant only in the late nonlinear stages of the simulation.
    Keywords: Aerodynamics
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    Direct Monte Carlo Simulations of Hypersonic Viscous Interactions Including Separation (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Results of calculations obtained using the direct simulation Monte Carlo method for Mach 25 flow over a control surface are presented. The numerical simulations are for a 35-deg compression ramp at a low-density wind-tunnel test condition. Calculations obtained using both two- and three-dimensional solutions are reviewed, and a qualitative comparison is made with the oil flow pictures highlight separation and three-dimensional flow structure.
    Keywords: Aerodynamics
    Type: NASA-TM-111528 , NAS 1.15:111528
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    Zonal Two Equation Kappa-Omega Turbulence Models for Aerodynamic Flows (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Two new versions of the kappa-omega two-equation turbulence model will be presented. The new Baseline (BSL) model is designed to give results similar to those of the original kappa-omega model of Wilcox, but without its strong dependency on arbitrary freestream values. The BSL model is identical to the Wilcox model in the inner 50% of the boundary-layer but changes gradually to the standard kappa-epsilon model (in a kappa- omega formulation) towards the boundary-layer edge. The free shear layers. The second version of the model is called Shear-Stress Transport (SST) model. It is a variation of the BSL model with the additional ability to account for the transport of the principal turbulent shear stress in adverse pressure gradient boundary-layers. The model is based on Bradshaw's assumption that the principal shear-stress is proportional to the turbulent kinetic energy, which is introduced into the definition of the eddy-viscosity. Both models are tested for a large number of different flowfields. The results of the BSL model are similar to those of the original kappa-omega model, but without the undesirable freestream dependency. The predictions of the SST model are also independent of the freestream values but show better agreement with experimental data for adverse pressure gradient boundary-layer flows.
    Keywords: Aerodynamics
    Type: NASA-TM-111629 , NAS 1.15:111629 , AIAA Paper 93-2906 , AIAA Fluid Dynamics Conference; Jul 06, 1993 - Jul 09, 1993; Orlando, FL; United States
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    Formulation for Simultaneous Aerodynamic Analysis and Design Optimization (1993)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: An efficient approach for simultaneous aerodynamic analysis and design optimization is presented. This approach does not require the performance of many flow analyses at each design optimization step, which can be an expensive procedure. Thus, this approach brings us one step closer to meeting the challenge of incorporating computational fluid dynamic codes into gradient-based optimization techniques for aerodynamic design. An adjoint-variable method is introduced to nullify the effect of the increased number of design variables in the problem formulation. The method has been successfully tested on one-dimensional nozzle flow problems, including a sample problem with a normal shock. Implementations of the above algorithm are also presented that incorporate Newton iterations to secure a high-quality flow solution at the end of the design process. Implementations with iterative flow solvers are possible and will be required for large, multidimensional flow problems.
    Keywords: Aerodynamics
    Type: NASA-CR-201036 , NAS 1.26:201036 , U.S. National Congress on Computational Mechanics; Aug 16, 1993 - Aug 18, 1993; Washington, DC; United States
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    Wing-Body Aeroelasticity Using Finite-Difference Fluid/Finite-Element Structural Equations on Parallel Computers (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a procedure for computing the aeroelasticity of wing-body configurations on multiple-instruction, multiple-data (MIMD) parallel computers. In this procedure, fluids are modeled using Euler equations discretized by a finite difference method, and structures are modeled using finite element equations. The procedure is designed in such a way that each discipline can be developed and maintained independently by using a domain decomposition approach. A parallel integration scheme is used to compute aeroelastic responses by solving the coupled fluid and structural equations concurrently while keeping modularity of each discipline. The present procedure is validated by computing the aeroelastic response of a wing and comparing with experiment. Aeroelastic computations are illustrated for a High Speed Civil Transport type wing-body configuration.
    Keywords: Aerodynamics
    Type: NASA-TM-111450 , NAS 1.15:111450 , AIAA Paper 94-1487 , AIAA/ASME/ASCE/AHS/ASC 35th Structural, Structural Dynamics, and Materials Conference; Apr 18, 1994 - Apr 20, 1994; Hilton Head, SC; United States
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    Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap (1993)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.
    Keywords: Aerodynamics
    Type: NASA-CR-203750 , NAS 1.26:203750 , AIAA Paper 93-3513 , Applied Aerodynamics; Aug 09, 1993 - Aug 11, 1993; Monterey, CA; United States
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    Electronic Resource
    Electronic Resource
    Flight of the honeybee (1992)
    Springer
    Journal of comparative physiology 162 (1992), S. 267-277 
    Details
    ISSN: 1432-136X
    Keywords: Aerodynamics ; Insect flight ; Body drag ; Drag coefficient ; Lift coefficient ; Honeybee
    Source: Springer Online Journal Archives 1860-2000
    Topics: Biology , Medicine
    Notes: Summary Drag forces and lift forces acting on honeybee trunks were measured by using specially built sensitive mechanical balances. Measurements were made on prepared bodies in ‘good’ and in ‘bad’ flight position, with and without legs, at velocities between 0.5 and 5m·s-1 (Reynolds numbers between 4·102 and 4·103) and at angles of attack between-20° and +20°. From the forces drag coefficients and lift coefficients were calculated. The drag coefficient measured with a zero angle of attack was 0.45 at 3≤v≤5m·s-1, 0.6 at 2m·s-1, 0.9 at 1m·s-1 and 1.35 at 0.5m·s-1, thus demonstrating a pronounced effect of Reynolds number on drag. These values are about 2 times lower (better) than those of a “drag disc” with the same diameter and attacked at the same velocity. The drag coefficient (related to constant minimal frontal area) was minimal at zero angle of attack, rising symmetrically to larger (+) and smaller (-) angles of attack in a non-linear fashion. The absolute value is higher and the rise is steeper at lower speeds or Reynolds numbers, but the incremental factors are independent of Reynolds number. For example, the drag coefficient is 1.44±0.05 times higher at an angle of attack of 20° than at one of 0°. On a double-logarithmic scale the slope of the drag versus Reynolds number plot was 1.5: with decreasing Reynolds number the relationship between drag and velocity changes from quadratic (Newton's law) to linear (viscous flow). Trunk drag was not systematically increased by the legs at any velocity or Reynolds number or any angle of attack. The legs appear to shape the trunk “aerodynamically”, to form a relatively low-drag trunk-leg system. The body is able to generate dynamic lift. Highly significant positive linear correlations between lift coefficient and angle of attack were determined for the trunk-leg system in the typical flight position. Lift coefficient was +0.05 at zero angle of attack (possibly attained during very fast flight), +0.1 at 5° (attained during fast flight), +0.25 at +20° (attained during slow flight) and +0.55 at 45° (attained whilst changing over to hovering). Average slope ΔcLΔα was 0.66±0.07, and average profile efficiency was 0.10. Non-wing lift contribution due to body form and banking only accounts for a few percent of body weight during fast flight. A non-wing lift contribution due to the legs has been demonstrated. The legs increase trunk lift by 23–24%. Reynolds number lift effects are present but of no biological significance. Force and power calculations do not support maximum flight speeds substantially higher than approximately 7m · s-1 relative to the ambient air. At this speed body drag attains 35% and body lift 8.4% of the body weight, and parasite power is 5% of the maximum metabolic power.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1007/BF00357534
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  • 57
    Electronic Resource
    Electronic Resource
    Euler solutions for self-generated rotor blade-vortex interactionsPresented as AIAA paper 90-1588 at the AIAA 20th Fluids and Plasma Dynamics Conference, Buffalo, New York, June 1990. (1992)
    Chichester : Wiley-Blackwell
    International Journal for Numerical Methods in Fluids 15 (1992), S. 427-451 
    Details
    ISSN: 0271-2091
    Keywords: Aerodynamics ; Rotor ; Blade-vortex ; Interactions ; Engineering ; Engineering General
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Mechanical Engineering, Materials Science, Production Engineering, Mining and Metallurgy, Traffic Engineering, Precision Mechanics
    Notes: A finite-difference procedure has been developed for the prediction of three-dimensional rotor blade-vortex interactions. The interaction velocity field was obtained through a non-linear superposition of the rotor flow field, computed using the unsteady three-dimensional Euler equations, and the embedded vortex wake flow field, computed using the law of Biot-Savart. In the Euler model, near wake rotational effects were simulated using the surface velocity ‘transpiration’ approach. As a result, a modified surface boundary condition was prescribed and enforced at each time step of the computations to satisfy the tangency boundary condition. For supercritical interactions using an upstream-generated vortex, accuracy of the numerical results were found to rely on the user-specified vortex core radius and vortex strength. For the more general self-generated subcritical interactions, vortex wake trajectories were computed using the lifting-line helicopter/rotor trim code CAMRAD. For these interactions, accuracy of the results were found to rely heavily on the CAMRAD-predicted vortex strength, vortex orientation with respect to the blade, and to a large extent on the user-specified vortex core radius. Results for the one-seventh scale model OLS rotor and for a non-lifting rectangular blade having a NACA0012 section are presented. Comparisons with the experimental windtunnel data are also made.
    Additional Material: 10 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/fld.1650150405
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    Sailboard Fin Design (1992)
    In:  CASI
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    Publication Date: 2018-06-09
    Description: In high performance boardsailing, demands on the vertical fin or "skeg" often produce "spinout" - when the skeg loses horizontal lift creating a force imbalance and causing the tail of the board to slide sideways. Richard Caldwell, RACE Technology, Inc. used NASA airfoil technology to solve this problem and formed a business based on his solution. After determining that the spinout resulted from air ventilating down the low pressure side of the underwater fin, he adapted the airfoil technology to the design of a short board skeg, which would overcome the problem and lower the drag, resulting in improved performance. He patented his RACE 145 foil section, formed his company and later returned to Langley for additional technical assistance. The company's newest product is a rigid sail that also incorporates NASA technology and has excellent performance. This company no longer exists - product is no longer in production.
    Keywords: Aerodynamics
    Type: Spinoff 1992; 60-61; NAA-NP-201
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    Characteristics of the Shuttle Orbiter Leeside Flow During A Reentry Condition (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A study of the leeside flow characteristics of the Shuttle Orbiter is presented for a reentry flight condition. The flow is computed using a point-implicit, finite-volume scheme known as the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). LAURA is a second-order accurate, laminar Navier-Stokes solver, incorporating finite-rate chemistry with a radiative equilibrium wall temperature distribution and finite-rate wall catalysis. The resulting computational solution is analyzed in terms of salient flow features and the surface quantities are compared with flight data.
    Keywords: Aerodynamics
    Type: AIAA Paper 92-2951
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    A Boom in Boomerangs (1992)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: Ted Bailey, a highly-ranked international boomerang designer and thrower, used information from a variety of NASA technical reports on aerodynamics and low-speed airfoils to design more competitive boomerangs. Because the boomerang is essentially an airfoil like an airplane wing, the technology transferred effectively and even contributed to the 1981 American victory over Australian throwers. In 1985, using four NASA reports, Bailey designed a new MTA (maximum time aloft) boomerang that broke the one-minute barrier, enabled throwers to throw and catch in less than three minutes and allowed competitors to complete the difficult "Super Catch" - five throw/catch sequences after launching the original boom while it is still aloft. Bailey is now considering other boomerang applications.
    Keywords: Aerodynamics
    Type: Spinoff 1992; 50-53; NASA-NP-201
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    Aerodynamic Spring and Damping of Free-Pitching Tips on a Semispan Wing (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A test was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel to derive aerodynamic spring and damping estimates for free-pitching tips on a semispan wing. The test model was a rectangular planform semispan wing with wing tips that had a single rigid-body pitch degree of freedom with respect to the inboard wing. A number of different tip planform geometries were tested, incorporating a range of quarter-chord sweep angles and taper ratios. The wing-tip dynamic response characteristics were measured at several wing angles of attack and tunnel dynamic pressures. The tip oscillations were initiated by releasing the tips from prescribed angles of attack. A new method to isolate Coulomb damping from aerodynamic damping from these tip-motion time histories is developed and applied. Correlations were performed between the experimentally derived wing-tip aerodynamic spring and damping values and predictions from a semiempirical analysis based on steady-state tip aerodynamic loads.
    Keywords: Aerodynamics
    Type: AIAA Paper 92-2111 , Dynamics Specialist Conference; Apr 16, 1992 - Apr 17, 1992; Dallas, TX; United States
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    Wing Download Results from a Test of a 0.658-Scale V-22 Rotor and Wing (1992)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A test of a 0.658-scale V-22 rotor and wing was conducted in the 40 x 80 Foot Wind Tunnel at Ames Research Center. One of the principal objectives of the test was to measure the wing download in hover for a variety of test configurations. The wing download and surface pressures were measured for a wide range of thrust coefficients, with five different flap angles, two nacelle angles, and both directions or rotor rotation. This paper presents these results, and describes a new method for interpreting wing surface pressure data in hover. This method shows that the wing flap can produce substantial lift loads in hover.
    Keywords: Aerodynamics
    Type: May 06, 1991 - May 08, 1991; Phoenix, AZ; United States|Journal of the American Helicopter Society; 58-63
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  • 63
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    Internal Shock Interactions in Propulsion/Airframe Integrated Three-Dimensional Sidewall Compression Scramjet Inlets (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The advantages and design requirements of propulsion/airframe integration for high Mach number flight have led to extensive study of the three-dimensional sidewall compression scramjet inlet in recent years. Recent research publications have indicated testing over a broad range of Mach number (2 to 18) in a variety of test gases, such as air, helium, and tetrafluoromethane. Multiple experimental techniques have been employed to obtain detailed internal shock interaction data, performance data, and inlet starting limits. Computational fluid dynamics has been effectively used for preliminary parametric studies as well as in parallel with experiments to aid in the explanation of unusual or unexpected flow phenomena. Inlets of this genre afford a relatively simple, generic geometry while producing a highly complex, three-dimensional flow field dominated by shock/shock and shock/boundary layer interactions. While the importance of the viscous effects in high speed inlet interactions is recognized, the present work addresses in a parametric fashion the inviscid effects of leading edge sweep, sidewall compression, and inflow Mach number on the internal shock structure in terms of inlet compression and mass capture. In the process, the source of the of the Mach number invariance with leading edge sweep for a constant sidewall compression class of inlet is identified, and a previously undocumented spillage phenomenon in a constant effective wedge angle class of inlets is discussed.
    Keywords: Aerodynamics
    Type: AIAA Paper 92-3099 , AIAA/SAE/ASME/ASEE 28th Joint Propulsion Conference and Exhibit; Jul 06, 1992 - Jul 08, 1992; Nashville, TN; United States
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  • 64
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    Direct Monte Carlo Simulations of Hypersonic Low-Density Flows about an ASTV Including Wake Structure (1992)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Results of a numerical study concerning flow past a 70-deg blunted cone in hypersonic low-density flow environments are presented using the direct simulation Monte-Carlo method. The flow conditions simulated are those that can be obtained in existing low-density hypersonic wind tunnels. Results indicate that a stable vortex forms in the near wake at and below a freestream Knudsen number (based on cone diameter) of 0.01 and the size of the vortex increases with decreasing Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases considered herein.
    Keywords: Aerodynamics
    Type: NASA-TM-111529 , NAS 1.15:111529
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  • 65
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    Design Software (1991)
    In:  CASI
    Details
    Publication Date: 2018-06-09
    Description: A NASA contractor and Small Business Innovation Research (SBIR) participant has converted its research into commercial software products for auto design, structural analysis and other applications. ViGYAN, Inc., utilizing the aeronautical research principle of computational fluid dynamics, has created - with VGRID3D and VPLOT3D - an easier alternative to conventional structured grids for fluid dynamic calculations.
    Keywords: Aerodynamics
    Type: Spinoff 1991; 114-115; NASA-NP-147
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  • 66
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    Study of Cluster Formation and its Effects on Rayleigh and Raman Scattering Measurements in a Mach 6 Wind Tunnel (1991)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Using a frequency-doubled Nd-YAG pulsed laser and a single-intensified CCD camera, Rayleigh scattering measurements have been performed to study the cluster formation in a Mach 6 wind tunnel at NASA Langley Research Center. These studies were conducted both in the free stream and in a model flow field for various flow conditions to gain an understanding of the dependence of the Rayleigh scattering (by clusters) on the local pressures and temperatures in the facility. Using the same laser system, we have also performed simultaneous measurements of the local temperature using the rotational Raman scattering of molecular nitrogen and determined the densities of molecular oxygen and nitrogen by using the vibrational Raman scattering from these species. Quantitative results will be presented in detail with emphasis on the applicability of the Rayleigh scattering for obtaining quantitative measurements of molecular densities both in the free stream and in the model flow field.
    Keywords: Aerodynamics
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  • 67
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    Topology of Flow Separation on Three-Dimensional Bodies (1991)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In recent years there has been extensive research on three-dimensional flow separation. There are two different approaches: the phenomenological approach and a mathematical approach using topology. These two approaches are reviewed briefly and the shortcomings of some of the past works are discussed. A comprehensive approach applicable to incompressible and compressible steady-state flows as well as incompressible unsteady flow is then presented. The approach is similar to earlier topological approaches to separation but is more complete and in some cases adds more emphasis to certain points than in the past. To assist in the classification of various types of flow, nomenclature is introduced to describe the skin-friction portraits on the surface. This method of classification is then demonstrated on several categories of flow to illustrate particular points as well as the diversity of flow separation. The categories include attached, two-dimensional separation and three different types of simple, three-dimensional primary separation, secondary separation, and compound separation. Hypothetical experiments are utilized to illustrate the topological terminology and its role in characterizing these flows. These hypothetical experiments use colored oil injected onto the surface at singular points in the skin-friction portrait. Actual flow-visualization information, if available, is used to corroborate the hypothetical examples.
    Keywords: Aerodynamics
    Type: NASA-CR-203748 , NAS 1.26:203748 , Applied Mechanics Reviews; 44; 7; 329-345
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  • 68
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    Secondary frequencies in the wake of a circular cylinder with vortex shedding (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: A detailed numerical study of two-dimensional flow past a circular cylinder at moderately low Reynolds numbers was conducted using three different numerical algorithms for solving the time-dependent compressible Navier-Stokes equations. It was found that if the algorithm and associated boundary conditions were consistent and stable, then the major features of the unsteady wake were well-predicted. However, it was also found that even stable and consistent boundary conditions could introduce additional periodic phenomena reminiscent of the type seen in previous wind-tunnel experiments. However, these additional frequencies were eliminated by formulating the boundary conditions in terms of the characteristic variables. An analysis based on a simplified model provides an explanation for this behavior.
    Keywords: Aerodynamics
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 225; 557-574
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  • 69
    Electronic Resource
    Electronic Resource
    Flight physiology of migrating Urania fulgens (Uraniidae) moths: kinematics and aerodynamics of natural free flight (1990)
    Springer
    Journal of comparative physiology 167 (1990), S. 145-154 
    Details
    ISSN: 1432-1351
    Keywords: Moth ; Flight ; Migration ; Kinematics ; Aerodynamics
    Source: Springer Online Journal Archives 1860-2000
    Topics: Biology , Medicine
    Notes: Summary Air speeds and wing kinematics were determined for the Neotropical moth Urania fulgens in natural migratory flight over Lake Gatun, Republic of Panama. Morphological parameters are presented for the same insects filmed in free flight. A quasi-steady aerodynamic analysis was used to show that unsteady mechanisms of lift generation are probably not necessary to produce the forces necessary for fast forward flight. Mechanical power requirements of forward flight were estimated from the biomechanical and morphological data. Over an airspeed range of 1.5 to 4.5 m/s, the mechanical power required to fly is predicted to increase dramatically with forward speed. A comparison of estimated metabolic rates with endogenous lipid reserves suggests that U. fulgens feeds extensively on flower nectar during migration from Central into South America.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1007/BF00192414
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  • 70
    Electronic Resource
    Electronic Resource
    The calculation of incompressible separated turbulent boundary layers (1990)
    Chichester : Wiley-Blackwell
    International Journal for Numerical Methods in Fluids 11 (1990), S. 39-56 
    Details
    ISSN: 0271-2091
    Keywords: Aerodynamics ; Turbulence ; Separation ; CLmax ; Laminar ; Bubble ; Engineering ; Engineering General
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Mechanical Engineering, Materials Science, Production Engineering, Mining and Metallurgy, Traffic Engineering, Precision Mechanics
    Notes: The algebraic turbulent model of Baldwin and Lomax was incorporated into the incompressible full Navier-Stokes code FIDAP. This model was extensively tested in the past in finite difference codes. We believe that the incorporation of the model also into the finite element code has resulted in a practical method to compute a variety of separated turbulent 2D flows. Firstly, we use the model to compute the attached flow about an aerofoil. Next, the application of the model to separated flows is presented by computing the flows at high angles of attack up to maximum lift. It is shown that the model is capable of predicting separation, steady stall and CLmax. As a difficult test of the model we compute the laminar separation bubble development directly using the full Navier-Stokes finite element code. As far as we know, this approach has not yet been reported. The importance of using an appropriate upwinding is discussed. When possible, comparison of computed results with experiments is presented and the agreement is good.
    Additional Material: 13 Ill.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1002/fld.1650110104
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  • 71
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    Secondary frequencies in the wake of a circular cylinder with vortex shedding (1990)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A detailed numerical study of two-dimensional flow past a circular cylinder at moderately low Reynolds numbers was conducted using three different numerical algorithms for solving the time-dependent compressible Navier-Stokes equations. It was found that if the algorithm and associated boundary conditions were consistent and stable, then the major features of the unsteady wake were well-predicted. However, it was also found that even stable and consistent boundary conditions could introduce additional periodic phenomena reminiscent of the type seen in previous wind-tunnel experiments. However, these additional frequencies were eliminated by formulating the boundary conditions in terms of the characteristic variables. An analysis based on a simplified model provides an explanation for this behavior.
    Keywords: Aerodynamics
    Type: NASA-CR-181998 , NAS 1.26:181998 , ICASE-90-16 , AD-A227099
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  • 72
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    The Importance of the Measurement of the Unsteady Wake of the Aeroassist Flight Experiment Vehicle: A White Paper (1990)
    In:  Other Sources
    Details
    Publication Date: 2019-07-10
    Description: The High Frequency Radiometer (HFR) is the only instrument on the Aeroassist Flight Experiment (AFE) with sufficient temporal resolution to discern the frequency of unsteady wake oscillations. Determining both the frequency and amplitude of wake unsteadiness during AFE atmospheric entry is essential for reliably predicting the geometry and motion of the wake of future Aeroassisted Space Transfer Vehicles (ASTV). These parameters directly affect the location and size of the pay- load and the weight of the required afterbody heat protection. The purpose of the AFE is to validate the technologies required for the design of ASTVs, which will be used to exploit Earth-lunar space. This validation will be conducted at a combination of vehicle size, altitude, and velocity not obtainable in ground-based facilities. The AFE will provide the experimental flight data needed to improve our understanding of hypersonic-wake physics and to validate computational predictions of the aerodynamic and heating loads, including afterbody radiative heating loads, on an ASTV. Reliable prediction of ASTV wake flows will ensure that payloads are located within the shear-layer envelope and will determine the amount of thermal protection the payloads require. Specifically, understanding the temporal nature of the wake unsteadiness is important for two reasons. Most importantly, analysis of ground-based experiments suggests that wake unsteadiness results in a variation of as much as +/- 5 deg in the shear-flow turning angle. This angle must be reliably predicted to avoid shear-layer impingement on the vehicle afterbody, which would result in heating rates of about 10 W/sq cm, of the same order as on the forebody stagnation point. Secondly, the energy associated with wake unsteadiness will reduce the static enthalpy of the wake fluid and cause an error of as much as 30% in the amount of predicted wake radiative heating. Therefore, the HFR flight data, which will quantify the frequency and amplitude of the wake unsteadiness, are required for the verification of computational models of ASTV flowfields. Measurement of wake unsteadiness has been endorsed by the Peer Science Steering Group and the AFE Computational Fluid Dynamics (CFD) Working Group. The Peer Science Steering Group has stressed that a reliable measurement of the wake unsteadiness is fundamental because incorporating unsteadiness represents a substantial challenge to the CFD community, and a reliable measurement will raise confidence in the computer simulation. The AFE CFD Working Group has stated that 44 any data which would prove the existence and effects of unsteady flow would be extremely valuable.
    Keywords: Aerodynamics
    Type: A-90321
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  • 73
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    Investigation of a 1/4-Scale Model of the Republic F-105 Airplane in the Langley 19-Foot Pressure Tunnel: Longitudinal Stability and Control of the Model Equipped with a Supersonic-Type Elliptical Wing-Root Inlet (1954)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: No abstract available
    Keywords: Aerodynamics
    Type: NACA-RM-SL54F28
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  • 74
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    Aerodynamic Characteristics at High and Low Subsonic Mach Numbers of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 Airfoil Sections at Angles of Attack from -2 Degrees to 30 Degrees (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation has been made in the Langley low-turbulence pressure tunnel of the aerodynamic characteristics of the NACA 0012, 64(sub 2)-015, and 64(sub 3)-018 airfoil sections. Data were obtained at Mach numbers from 0.3 to that for tunnel choke, at angles of attack from -2deg to 30deg, and with the surface. of each airfoil smooth-and with roughness applied at the leading edge.The Reynolds numbers of the tests ranged from 0.8 x 10(exp 6) to 4.4 x 10(exp 6). The results are presented as variations of lift, drag, and quarter-chord pitching-moment coefficients with Mach number.
    Keywords: Aerodynamics
    Type: NACA-RM-L54H06a
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  • 75
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    Effect of a Leading-edge Flap upon the Lift, Drag, and Pitching Moment of an Airplane Employing a Thin, Unswept Wing (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: 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.
    Keywords: Aerodynamics
    Type: NACA-RM-SA54B16
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  • 76
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    Preliminary Altitude Performance Characteristics of the J57-P-1 Turbojet Engine with Fixed-area Exhaust Nozzle (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation to determine the altitude performance of the J57-P-1 turbojet engine and components was conducted at the NACA Lewis altitude wind tunnel. Data were obtained over a corrected inboard rotor speed range from 56 to 106 percent of rated speed, with intercompressor bleeds both open and closed, at altitudes from 15,000 to 50,000 feet and at a flight Mach number of 0.81. The corresponding range of Reynolds number indices was from 0.858 to 0.213. All data presented were obtained with a fixed-area exhaust nozzle sized according to the manufacturer's specification. Over-all engine performance parameters are presented as functions of inboard rotor speed corrected on the basis of engine inlet temperature. Component parameters are presented as functions of their respective corrected rotor speeds. A tabulation of all performance data is included in addition to the graphical presentation. Corrected net thrust is unusually sensitive to changes in corrected inboard rotor speed in the high speed region. A change of 1 percent in speed, at sated speed, produced a change of 6 percent in corrected net thrust . At rated engine speed, increasing the altitude from 15,000 to 50,000 feet at a constant flight Mach number of 0.81 increased the specific fuel consumption 13 percent but did not affect corrected net thrust.
    Keywords: Aerodynamics
    Type: NACA-RM-SE54D30
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  • 77
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    Effect on Drag of Longitudinal Positioning of Half-Submerged and Pylon-Mounted Douglas Aircraft Stores on a Fuselage with and without Cavities between Mach Numbers 0.9 and 1.8 (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The effect on drag of positioning symmetrically mounted Douglas Aircraft Company, Inc. stores in pairs on a parabolic fuselage of fineness ratio 10.0 has been determined by flight tests of rocket-propelled, zero-lift models through a range of Mach number from 0.9 to 1.8. The stores were mounted in half-submerged positions and on pylons and were tested in three longitudinal locations on the fuselage with the forward position being located at the maximum diameter of the fuselage. The effects on drag of removing the half-submerged stores or extending them outward on pylons also was investigated by tests of models with half-submerged-store cavities on the fuselage. Two pylons differing in airfoil section and thickness were tested at the forward position of the stores on the fuselage with cavities. The half-submerged stores gave the smallest drag increments, which were approximately equal regardless of their respective longitudinal locations. Removing the half-submerged stores to expose the cavities increased the drag increments from two to three times. For the pylon-mounted stores, the store in the midposition had less drag than in the forward or rear positions at supersonic speeds. Adding the half-submerged-store cavities to the pylon-mounted-store configurations reduced the drag at the rear position between Mach numbers 0.95 and 1.50 and increased the drag at the midposition throughout the speed range. Changing from the 6-percent-thick flat pylon to the 10-percent-thick airfoil pylon increased the total drag slightly above Mach number 1.10. Good agreement was obtain& between the experimental and theoretical interference drag coefficients for the pylon-mounted stores (without fuselage cavities} in the three longitudinal locations tested at Mach numbers 1.2 and 1.5.
    Keywords: Aerodynamics
    Type: NACA-RM-L54E26
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  • 78
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    Longitudinal Aerodynamic Characteristics to Large Angles of Attack of a Cruciform Missile Configuration at a Mach Number of 2 (1954)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: The lift, pitching-moment, and drag characteristics of a missile configuration having a body of fineness ratio 9.33 and a cruciform triangular wing and tail of aspect ratio 4 were measured at a Mach number of 1.99 and a Reynolds number of 6.0 million, based on the body length. The tests were performed through an angle-of-attack range of -5 deg to 28 deg to investigate the effects on the aerodynamic characteristics of roll angle, wing-tail interdigitation, wing deflection, and interference among the components (body, wing, and tail). Theoretical lift and moment characteristics of the configuration and its components were calculated by the use of existing theoretical methods which have been modified for application to high angles of attack, and these characteristics are compared with experiment. The lift and drag characteristics of all combinations of the body, wing, and tail were independent of roll angle throughout the angle-of-attack range. The pitching-moment characteristics of the body-wing and body-wing-tail combinations, however, were influenced significantly by the roll angle at large angles of attack (greater than 10 deg). A roll from 0 deg (one pair of wing panels horizontal) to 45 deg caused a forward shift in the center of pressure which was of the same magnitude for both of these combinations, indicating that this shift originated from body-wing interference effects. A favorable lift-interference effect (lift of the combination greater than the sum of the lifts of the components) and a rearward shift in the center of pressure from a position corresponding to that for the components occurred at small angles of attack when the body was combined with either the exposed wing or tail surfaces. These lift and center-of-pressure interference effects were gradually reduced to zero as the angle of attack was increased to large values. The effect of wing-tail interference, which influenced primarily the pitching-moment characteristics, is dependent on the distance between the wing trailing vortex wake and the tail surfaces and thus was a function of angle of attack, angle of roll, and wing-tail interdigitation. Although the configuration at zero roll with the wing and tail in line exhibited the least center-of-pressure travel, the configuration with the wing and tail interdigitated had the least change in wing-tail interference over the angle-of-attack range. The lift effectiveness of the variable-incidence wing was reduced by more than 70 percent as a result of an increase in the combined angle of attack and wing incidence from 0 deg to 40 deg. The wing-tail interference (effective downwash at the tail) due to wing deflection was nearly zero as a result of a region of negative vorticity shed from the inboard portion of the wing. The lift characteristics of the configuration and its components were satisfactorily predicted by the calculated results, but the pitching moments at large angles of attack were not because of the influence of factors for which no adequate theory is available, such as the variation of the crossflow drag coefficient along the body and the effect of the wing downwash field on the afterbody loading.
    Keywords: Aerodynamics
    Type: NACA-RM-A54H27
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  • 79
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    Flight Determination of the Longitudinal Stability Characteristics of a 0.133-Scale Rocket-Powered Model of the Consolidated Vultee XFY-1 Airplane without Propellers at Mach Numbers from 0.73 to 1.19, TED No. NACA DE 369 (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A flight test has been conducted to determine the longitudinal stability and control,characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane without propellers for the Mach number range between 0.73 and 1.19.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54B03A
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  • 80
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    Wind-Tunnel Investigation of the Effects on the Aerodynamic Characteristics of Modifications to a Model of a Bomb Mounted on a Wing-Fuselage Model and to a Model of the Bomb Alone (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel to determine effects of modifications to a bomb model (particularly with regard to drag) when mounted on a wing-fuselage model and tested at Mach numbers from 0.70 to 1.10. In addition, the static longitudinal stability characteristics of several configurations of a larger scale model of the bomb alone were obtained over a Mach number range from 0.50 to 0.95. The results obtained for the wing-fuselage-bomb model indicate that large reductions in installation drag were obtained for the wing-fuselage-bomb model when the flat nose of the basic bomb was replaced by rounded or pointed noses of various calibers. Shortening the mounting pylon gave further decreases in the installation drag. The tests of the bomb alone indicated that only the flat-nose configurations were stable over the greater part of the Mach number range. Nose-shape modifications which improved the drag also caused the bombs to become unstable at low angles of attack. The stability of the low-drag bomb configurations could be improved by lengthening the cylindrical portion of the body behind the center of gravity.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54D30
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  • 81
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    Effects of Fuselage Modifications on the Drag Characteristics of a 1/20-Scale Model of the Convair F-102 Airplane at Transonic Speeds (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation has been conducted in the Langley 8-foot transonic tunnel to determine the effects of several fuselage modifications on the transonic drag-rise characteristics of a 1/20-scale model of the Convair F-102 airplane. Tests covered an angle-of-attack range from 0deg to about 10deg and a Mach number range from 0.60 to 1.14. Results indicated that the transonic drag rise .for the basic F-102 airplane could be substantially reduced by extending the fuselage after-body approximately 8 percent of the fuselage length. Tests of other bodies indicated that a shorter (4-percent) afterbody extension may have a similar effect on the drag rise. Further improvement of the axial cross-sectional-area distribution of the 8-percent extended configuration through the addition of fuselage volume resulted in additional reductions in the drag rise at a Mach number of 1.0 and caused no or only slight drag penalties at the higher Mach numbers. The results of the present tests generally substantiate the area-rule concept with respect to the prediction of the transonic drag rise through the use of an equivalent-area body of revolution for a practical delta-wing airplane configuration.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54K18a
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  • 82
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    Preliminary Low-Speed Wind-Tunnel Investigation of Some Aspects of the Aerodynamic Problems Associated with Missiles Carried Externally in Positions Near Airplane Wings (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A low-speed wind-tunnel investigation has been made of some aspects of the aerodynamic problems associated with the use of air-to-air missiles when carried externally on aircraft. Measurements of the forces and moments on a missile model for a range of positions under the mid-semispan location of a 45deg sweptback wing indicated longitudinal and lateral forces with regard to both carriage and release of the missiles. Surveys of the characteristics of the flow field in the region likely to be traversed by the missiles showed abrupt gradients in both flow angularity and in local dynamic pressure. Through the use of aerodynamic data on the isolated missile and the measured flow-field characteristics, the longitudinal forces and moments acting on the missile while in the presence of the wing-fuselage combination could be estimated with fair accuracy. Although the lateral forces and moments predicted were qualitatively correct, there existed some large discrepancies in absolute magnitude.
    Keywords: Aerodynamics
    Type: NACA-RM-L54J20
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  • 83
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    Wind-Tunnel Investigation at Subsonic and Supersonic Speeds of a Fighter Model Employing a Low-Aspect-Ratio Unswept Wing and a Horizontal Tail Mounted Well Above the Wing Plane - Longitudinal Stability and Control (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Experimental results showing the static longitudinal-stability and control characteristics of a model of a fighter airplane employing a low-aspect-ratio unswept wing and an all-movable horizontal tail are presented. The investigation was made over a Mach number range from 0.60 to 0.90 and from 1.35 to 1.90 at a constant Reynolds number of 2.40 million, based on the wing mean aerodynamic chord. Because of the location of the horizontal tail at the tip of the vertical tail, interference was noted between the vertical tail and the horizontal tail and between the wing and the horizontal tail. This interference produced a positive pitching-moment coefficient at zero lift throughout the Mach number range of the tests, reduced the change in stability with increasing lift coefficient of the wing at moderate lift coefficients in the subsonic speed range, and reduced the stability at low lift coefficients at high supersonic speeds. The lift and pitching-moment effectiveness of the all movable tail was unaffected by the interference effects and was constant throughout the lift-coefficient range of the tests at each Mach number except 1.90.
    Keywords: Aerodynamics
    Type: NACA-RM-SA54D05
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  • 84
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    Static Longitudinal and Lateral Stability and Control Characteristics of a 1/15-Scale Model of the Grumman F9F-9 Airplane at a Mach Number of 1.41 (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.41 to determine the static stability and control and drag characteristics of a l/l5-scale model of the Grunman F9F-9 airplane. The effects of alternate fuselage shapes, wing camber, wing fences, and fuselage dive brakes on the aerodynamic characteristics were also investigated. These tests were made at a Reynolds number of 1.96 x l0 (exp 6) based on the wing mean aerodynamic chord of 0.545 foot. The basic configuration had a static margin of stability of 38.4 percent of the mean aerodynamic chord and a minimum drag coefficient of 0.049. For the maximum horizontal tail deflection investigated (-l0 deg), the maximum trim lift coefficient was 0.338. The basic configuration had positive static lateral stability at zero angle of attack and positive directional control throughout the angle-of-attack range investigated up to ll deg.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54G08
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  • 85
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    Longitudinal Stability and Control Characteristics at Transonic Speeds of a 1/30-Scale Model of the Republic XF-103 Airplane (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The longitudinal stability and control characteristics of a 1/30-scale model of the Republic XF-103 airplane were investigated in the Langley 8-foot transonic tunnel. The effect of speed brakes located at the end of the fuselage was also investigated. The main part of the investigation was made with internal flow in the model, but some data were obtained with no internal flow. The longitudinal stability and control at transonic-speeds appeared satisfactory. The transonic drag rise was small. The speed brakes had no adverse effects on longitudinal stability.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54H24
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  • 86
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    Free-Flight Zero-Lift Drag Results from a 1/5-Scale Model and Several Small-Scale Equivalent Bodies of Revolution of the Convair F-102 Configuration at Mach Numbers up to 1.34 (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun. The several small-scale models tested corresponded to: the basic configuration, the 1/5-scale rocket-propelled model configuration, a 2-foot (full-scale) fuselage-extension configuration, and a 7-foot (full-scale) fuselage-extension configuration. Models designed to correspond to the area distribution at a Mach number of 1.0 were flown for each of these 'shapes and, in addition, models designed to correspond to the area distribution at a Mach number of 1.2 were flown for the 1/5-scale rocket-propelled model and the 7-foot-fuselage-extension configuration. The value of external pressure drag coefficient (including base drag) obtained from the large-scale rocket model was 0.0190 at a Mach number of 1..05 and the corresponding values from the equivalent-body tests varied from 0.0183 for the rocket-propelled model shape to 0.0137 for the 7-foot-fuselage-extension configuration. From the results of tests of equivalent bodies designed to correspond to the area distribution at a Mach number of 1.0, it is evident that the small changes in shape incorporated in the basic and 2-foot-fuselage-extension configurations from that of the rocket-propelled model configuration will provide no significant change in pressure drag. On the other hand, the data from the 7-foot-fuselage-extension model indicate a substantial reduction in pressure drag at transonic speeds.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54DO9b
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  • 87
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    A Transonic Wind-Tunnel Investigation of the Longitudinal Aerodynamic Characteristics of a Model of the Lockheed XF-104 Airplane (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The transonic longitudinal aerodynamic characteristics of a 0.0858-scale model of the Lockheed XF-104 airplane have been obtained from tests at the Langley 16-foot transonic tunnel. The results of the investigation provide some general information applicable to the transonic properties of thin, low-aspect-ratio, unswept wing configurations utilizing a high horizontal tail . The model employs a horizontal tail mounted at the top of the vertical tail and a wing with an aspect ratio of 2.5, a taper ratio of 0.385, and 3.4-percent-thick airfoil sections. The lift, drag, and static longitudinal pitching moment were measured at Mach numbers from 0.80 t o 1.09 and angles of attack from -2.5 deg to 22.5 deg. Some of the dynamic longitudinal stability properties of the airplane have been predicted from the test results. In addition, some visual flow studies on the wing surfaces obtained at Mach numbers of 0.80 and 1.00 are included. Results of the investigation show that the transonic rise in drag coefficient at zero lift is about 0.030. At high angles of attack, the model becomes longitudinally unstable at Mach numbers from 0.80 t o 0.90, whereas a reduction in static stability is experienced when very high angles of attack are reached at Mach numbers above 0.90. Longitudinal dynamic stability calculations show that the longitudinal control is good at angles of attack below the unstable break in the static pitching-moment curves, but a typical corrective control applied after the occurrence of neutral stability has little effect in averting pitch-up.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54K19a
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  • 88
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    Longitudinal Control Characteristics of a 1/20-Scale Model of the Convair F-102 Airplane at Transonic Speeds (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The effects of elevator deflections from 0deg to -20deg on the force and moment characteristics of a 1/20-scale model of the Convair F-102 airplane with chordwise fences have been determined a t Mach numbers from 0.6 to 1.1 for angles of attack up to 20deg in the Langley 8-foot transonic tunnel. The configuration exhibited static longitudinal stability throughout the range tested, although a mild pitch-up tendency was indicated a t Mach numbers from 0.85 to 0.95. Elevator pitch effectiveness decreased rapidly between the Mach numbers of 0.9 and 1.0, however, no complete loss or reversal was indicated for all conditions tested. Because of the type of longitudinal control used, trimming the configuration from the zero elevator condition resulted in substantial decreases in lift-curve slope and maximum lift-drag ratio and increases in drag due to lift. The drag at zero lift, drag due to lift, and trim drag were high for this configuration.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54G15
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  • 89
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    Flight Results from a 1/10-Scale Rocket Model of the Lockheed XF-104 Airplane at Transonic Mach Numbers (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A 1/10-scale rocket model of the Lockheed XF-104 with faired inlets has been flown over a Mach number range from 0.80 to 1.45 to determine low-lift drag and a limited amount of stability data. The center-of-gravity locations were 4.0 and 1.5 percent of the mean aerodynamic chord before and after sustainer firing, respectively. Oscillations induced by pulse rockets were used to determine stability data. The external transonic drag coefficient increased from a value of 0.0160 at Mach number 0.80 to a maximum of 0.0432 near Mach number 1-13, with a drag rise Mach number of about 0.93. At Mach numbers where it could be determined, the model exhibited stable dynamic and static stability characteristics at low lift.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54E14
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  • 90
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    Longitudinal Stability Characteristics of the Consolidated Vultee XFY-1 Airplane with Windmilling Propellers as Obtained from Flight of 0.133-Scale Rocket-Propelled Model at Mach Numbers from 0.70 to 1.13 (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A flight test has been conducted to determine the longitudinal stability and control characteristics of a 0.133-scale model of the Consolidated Vultee XFY-1 airplane with windmilling propellers for the Mach number range between 0.70 and 1.13. The variation of lift-curve slope C(sub L(sub alpha) with Mach number was gradual with a maximum value of 0.074 occurring at a Mach number of 0.97. Propellers had little effect upon the values of lift-curve slope or the linearity of lift coefficient with angle of attack. At lift coefficients between approximately 0.25 and 0.45 with an elevon angle of approximately -l0 deg, there was a region of neutral longitudinal stability at Mach numbers below 0.93 introduced by the addition of windmilling propellers. Below a lift coefficient of 0.10 and above a lift coefficient of 0.45, the model was longitudinally stable throughout the Mach number range of the test. There was a forward shift in the aerodynamic center of about 3-percent mean aerodynamic chord introduced by the addition of propellers. The aerodynamic center as determined at low lift moved gradually from a value of 28.5-percent mean aerodynamic chord at a Mach number of 0.75 to a value of 47-percent mean aerodynamic chord at a Mach number of 1.10. There was an abrupt decrease in pitch damping between Mach numbers of 0.88 and 0.99 followed by a rapid increase in damping to a Mach number of 1.06. The propellers had little effect upon the pitch damping characteristics . The transonic trim change was a large pitching-down tendency with and without windmilling propellers. The elevons were effective pitch controls throughout the speed range; however, their effectiveness was reduced about 50 percent at supersonic speeds. The propellers had no appreciable effect upon the control effectiveness.
    Keywords: Aerodynamics
    Type: NACA-RM-SL54F11
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  • 91
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    Performance Characteristics of Plane-Wall Two-Dimensional Diffusers (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experiments have been made at Stanford University to determine the performance characteristics of plane-wall, two-dimensional diffusers which were so proportioned as to insure reasonable approximation of two-dimensional flow. All of the diffusers had identical entrance cross sections and discharged directly into a large plenum chamber; the test program included wide variations of divergence angle and length. During all tests a dynamic pressure of 60 pounds per square foOt was maintained at the diffuser entrance and the boundary layer there was thin and fully turbulent. The most interesting flow characteristics observed were the occasional appearance of steady, unseparated, asymmetric flow - which was correlated with the boundary-layer coalescence - and the rapid deterioration of flow steadiness - which occurred as soon as the divergence angle for maximum static pressure recovery was exceeded. Pressure efficiency was found to be controlled almost exclusively by divergence angle, whereas static pressure recovery was markedly influenced by area ratio (or length) as well as divergence angle. Volumetric efficiency. diminished as area ratio increased, and at a greater rate with small lengths than with large ones. Large values of the static-pressure-recovery coefficient were attained only with long diffusers of large area ratio; under these conditions pressure efficiency was high and. volumetric efficiency low. Auxiliary tests with asymmetric diffusers demonstrated that longitudinal pressure gradient, rather than wall divergence angle, controlled flow separation. Others showed that the addition of even a short exit duct of uniform section augmented pressure recovery. Finally, it was found that the installation of a thin, central, longitudinal partition suppressed flow separation in short diffusers and thereby improved pressure recovery
    Keywords: Aerodynamics
    Type: NACA-TN-2888
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  • 92
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    Theoretical Performance Characteristics of Sharp Lip Inlets at Subsonic Speeds (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method is presented for the estimation of the subsonic-flight-speed characteristics of sharp-lip inlets applicable to supersonic aircraft. The analysis, based on a simple momentum balance consideration, permits the computation of inlet pressure recovery - mass-flow relations and additive-drag coefficients for forward velocities from zero to the speed of sound. The penalties for operation of a sharp-lip inlet at velocity ratios other than 1.0 may be severe; at lower velocity ratios an additive drag is incurred that is not cancelled by lip suction, while at higher velocity ratios, unavoidable losses in inlet total pressure will result. In particular, at the take-off condition, the total pressure and the mass flow for a choked inlet are only 79 percent of the values ideally attainable with a rounded lip. Experimental data obtained at zero speed with a sharp-lip supersonic inlet model were in substantial agreement with the theoretical results.
    Keywords: Aerodynamics
    Type: NACA-TN-3004
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  • 93
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    On the Development of Turbulent Wakes from Vortex Streets (1953)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wake development behind circular cylinders at Reynolds numbers from 40 to 10,000 was investigated in a low-speed wind tunnel. Standard hotwire techniques were used to study the velocity fluctuations. The Reynolds number range of periodic vortex shedding is divided into two distinct subranges. At R = 40 to 150, called the stable range, regular vortex streets are formed and no turbulent motion is developed. The range R = 150 to 300 is a transition range to a regime called the irregular range, in which turbulent velocity fluctuations accompany the periodic formation of vortices. The turbulence is initiated by laminar-turbulent transition in the free layers which spring from the separation points on the cylinder. This transition first occurs in the range R = 150 to 300. Spectrum and statistical measurements were made to study the velocity fluctuations. In the stable range the vortices decay by viscous diffusion. In the irregular range the diffusion is turbulent and the wake becomes fully turbulent in 40 to 50 diameters downstream. It was found that in the stable range the vortex street has a periodic spanwise structure. The dependence of shedding frequency on velocity was successfully used to measure flow velocity. Measurements in the wake of a ring showed that an annular vortex street is developed.
    Keywords: Aerodynamics
    Type: NACA-TN-2913
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  • 94
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    The Base Pressure at Supersonic Speeds on Two-Dimensional Airfoils and Bodies of Revolution (With and Without Fins) Having Turbulent Boundary Layers (1953)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An analysis has been made of available experimental data to show the effects of most variables that are predominant in determining base pressure at supersonic speeds. Two dimensional bases and bases of bodies of revolution, restricted to turbulent boundary layers, are covered.
    Keywords: Aerodynamics
    Type: NACA-RM-L53C02
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  • 95
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    Elastic Buckling under Combined Stresses of Flat Plates with Integral Waffle-Like Stiffening (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Theory and experiment were compared and found in good agreement for the elastic Buckling under combined stresses of long flat plates with integral waffle-like stiffening in a variety of configurations. For such flat plates, 45deg waffle stiffening was found to be the most effective of the configurations for the proportions considered over the widest range of combinations of compression and shear.
    Keywords: Aerodynamics
    Type: NACA-RM-L53J27
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  • 96
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    Rocket-Model Measurements of Zero-Lift Damping in Roll of the Bell MX-776 Missile at Mach Numbers from 0.6 to 1.56 (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The zero-lift damping in roll of the Bell MX-776 missile has been measured by a sting-mounted rocket-model technique at Mach numbers from 0.6 to 1.56. The damping-in-roll data, in general, show no unusual variation with Mach number. Aileron rolling-moment effectiveness derived from these data and previously obtained rolling-effectiveness data appear reasonable,
    Keywords: Aerodynamics
    Type: NACA-RM-SL54A13
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  • 97
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    Lateral Oscillatory Characteristics of the Republic F-91 Airplane Calculated by Using Low-Speed Experimental Static and Rotary Derivatives (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The present investigation was conducted to determine, from low-speed tests in the Langley stability tunnel, the static and rotary derivatives of a 1/9-scale model of the Republic F-91 airplane and various of its components (including the effects of wing incidence) and to determine the accuracy with which the period and damping of the lateral oscillation of the airplane could be calculated by using these experimentally between flight and calculated period and damping of the lateral oscillation were made for Mach numbers from 0.4 to 0.9 at an altitude of 20,OOO feet for 0deg wing incidence and several other wing incidences. Some comparisons were made of the static and rotary derivatives of the model and derivatives estimated by available procedures. determined derivatives (corrected for Mach number effects). Comparisons The results of the investigation have indicated that the model did not have unusual aerodynamic characteristics except for a large (about -0.125) increment in the damping in yaw contributed by the fuselage. Changes in wing incidence, in general, had little effect on the static and rotary derivatives of the model. The static and rotary derivatives of the model could be estimated with good accuracy only in the low angle-of-attack range by using available procedures.
    Keywords: Aerodynamics
    Type: NACA-RM-L53G01
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  • 98
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    An Exploratory Investigation of Some Types of Aeroelastic Instability of Open and Closed Bodies of Revolution Mounted on Slender Struts (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Aeroelastic instability phenomena of isolated open and closed rigid bodies of revolution free to move under elastic restraint have been investigated experimentally at low speeds by means of models suspended at zero angles of attack and yaw on slender flexible struts from a wind tunnel ceiling. Three types of instability were observed - flutter similar to classical bending-torsion flutter, divergence, and an uncoupled oscillatory instability which consists in nonviolent continuous or intermittent small-amplitude oscillations involving only angular deformations. The speeds at which this oscillatory instability starts were found to be as low as about one-third of the speed at flutter or divergence and to depend on the shape of the body, particularly that of the afterbody, and on the relative location of the elastic axis. An attempt has been made to calculate the airspeeds and, in the case of the oscillatory phenomena, the frequencies at which these instabilities occur by using slender-body theory for the aerodynamic forces on the bodies and neglecting the aerodynamic forces on the struts. However, the agreement between the speeds and frequencies calculated in this manner and those actually observed has been found to be generally unsatisfactory; with the exception of the frequencies of the uncoupled oscillations which could be predicted with fair accuracy. The nature of the observed phenomena and of the forces on bodies of revolution suggests that a significant improvement in the accuracy of analytical predictions of these aeroelastic instabilities can be had only by taking into account the effects of boundary-layer separation on the aerodynamic forces.
    Keywords: Aerodynamics
    Type: NACA-RM-L53E07
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  • 99
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    Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.
    Keywords: Aerodynamics
    Type: NACA-RM-L53I10
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  • 100
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    Investigation of the Attenuation of Plane Shock Waves Moving over very Rough Surfaces (1953)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Experimental measurements of the attenuation of plane shock waves moving over rough walls have been made in a shock tube. Measurements of the boundary-layer characteristics, including thickness and velocity distribution behind the shock, have also been made with the aid of new cal techniques which provide direct information on the local boundary-layer conditions at the rough walls. Measurements of shock speed and shock pressure ratio are presented for both smooth-wall and rough-wall flow over lengths of machined-smooth and rough strips which lined all four walls of the shock tube. A simplified theory based on Von Karman's expression for skin-friction coefficient for flow over rough walls, along with a wave-model concept and extensions to include time effects, is presented. In this theory, the shock-tube flow is assumed to be one-dimensional at all times and the wave-model concept is used to relate the local layer growth to decreases in shock strength. This concept assumes that local boundary-layer growths act as local mass-flow sinks, which give rise to expansion waves which, in turn, overtake the shock and lower its mass flow accordingly.
    Keywords: Aerodynamics
    Type: NACA-RM-SL53D13A
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