Scientific and Technical Aerospace Reports A Semimonthly Publication of the National Aeronautics and Space Administration Sponsored by AFIT, Virginia L. Sullivan 1990 131 P The B-2 Stealth bomber was labelled by some media as the most expensive airplane ever, while others have proclaimed it as the most innovative and most efficient aircraft ever built. This dichotomy of the press can be attributed to the sources being quoted in a specific article, according to the results of this study. The content analysis of hard news stories in the New York Times, the Washington Post, and the Los Angeles Times revealed that from the 13 categories analyzed as favorable or unfavorable, 55 percent of the 365 articles were favorable; the remainder was classified neutral because the direction equalled zero. The qualitative part of this study consisted of personal and telephone interviews with military and political leaders. Although some military leaders expressed contempt for the media, they mainly pointed fingers at Congress for being tenuous in statements about the B-2 to the press. Since certain aspects of the media coverage were unfavorable, the media appeared to rely more on the agenda set by Congress than the military leaders' agenda. However, since both Congress and the military placed the B-2 prominently on their agendas, so did the press. GRA conventional pressure sensing devices do not provide meaningful measurements. However, a hot filament gauge was developed and miniaturized which will measure the pressure in the 10(exp -1) to 10(exp -5) torr (2 millipsi to 0.2 micropsi) region, hence the name Micropsi gauge. Laboratory studies were made comparing the currently available devices with the newly developed miniature low power 'Micropsi' pressure sensor. Author N91-11668 Georgia Inst. of Tech., Atlanta. A NOVEL APPROACH FOR ANALYZING SUPERSONIC HIGH REYNOLDS NUMBER FLOWS WITH SEPARATION Ph.D. Thesis Gregory Dan Power 1990 179 p Avail: Univ. Microfilms Order No. DA9023826 A physically based computational technique to solve the steady Navier-Stokes equations for the above class of problems is developed. This is accomplished by tailoring the analysis to reflect the dominant mechanism of strong viscous/inviscid interaction as determined from experimental observations and mathematical theories. A reduced set of composite equations, derived from the full Navier-Stokes equations, is solved in a globally-iterated, space-marching manner where upstream influence is modelled through the introduction of the stream function along a single line oriented in the streamwise direction and located within the viscous region. This technique for solving the composite equations represents an extension of established two-layer interacting viscous-layer approaches. It should be applicable to a broader range of flow conditions than the two-layer methods while providing a comparable level of efficiency and thus could provide an effective alternative to time-marching algorithms. Computational results are presented for the cases of normal slot injection into a supersonic stream and shock/boundary-layer interaction with and without separation. These results compare favorably with results computed using an asymptotic analysis, results computed using a time-marching Navier-Stokes procedure, and available experimental data. In addition to verifying that the current technique accurately models complex supersonic flows with strong viscous/inviscid interaction, the results also confirm that for steady flows the effects of the inviscid elliptic characteristics associated with the small subsonic region near the surface are negligibly small compared to the effects of the displacement-thickness/pressure-gradient interaction mechanism. Dissert. Abstr. N91-11669 Southampton Univ. (England). Timothy Peter Roberts 1989 227 p The results of an investigation into the static and dynamic behavior of the hypersonic separated flow generated over a flat plate-rearward flap configuration are reported with particular emphasis on the dynamic aspects. A range of flap deflection angles up to 35 degrees were considered. For the dynamic tests, the form of the motion of the flap was such that it was rapidly driven up from an almost zero degree angle of attack, to one of about 40 degrees, and then rapidly reversed. The investigation involved some theoretical modelling, and an experimental program in which extensive static pressure measurements and flow visualization studies were made. The experimental results showed that the rapid motion of the flap introduced some unsteadiness into the flow over the model, particularly near reattachment. In addition it was observed that the growth or decay of the separation region 'lags' behind the changing flap angle, and that the extent of this lag is to first order, proportional to the flap angular velocity. Using this linear relationship, the experimental data were interpreted to obtain the difference in flap angles at which a given position of separation would be obtained, between that for the flap rising and that for the flap falling, both with a flap angular velocity of 51 rad/s. The experimental results have shown that the overall force and moment coefficients on the model, for the flap angular velocity range of -51 to +51 rad/s, differ by no more than -3.8 percent of the quasi-steady value. Dissert. Abstr. N91-11670 Notre Dame Univ., IN. COMPRESSIBLE FLOWS WITH PERIODIC VORTICAL DISTURBANCES AROUND LIFTING AIRFOILS Ph.D. Thesis James Russell Scott 1990 232 p Avail: Univ. Microfilms Order No. DA9023948 A numerical method is developed for solving periodic, three-dimensional, vortical flows around lifting airfoils in subsonic flow. This first-order method fully accounts for the distortion effects of the nonuniform mean flow on the convected upstream vortical disturbances. The unsteady velocity is split into a vortical component and an irrotational field whose potential satisfies a nonconstant-coefficient, inhomogeneous, and convective wave equation. Using an elliptic coordinate transformation, the unsteady boundary value problem is solved in the frequency domain on grids. In general, the agreement between the numerical and analytical results is very good for reduced frequencies ranging from 0 to 4, and for Mach numbers ranging from .1 to .8. Numerical results are also presented for a wide variety of flow configurations to determine the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. Each of these parameters can have a significant effect on the unsteady airfoil response to the incident disturbances, and the effect depends strongly upon the reduced frequency and the dimensionality of the gust. For a one-dimensional or two-dimensional gust, the results indicate that airfoil thickness increases the unsteady lift and moment at the low reduced frequencies but decreases it at the high reduced frequencies. It is shown that mean airfoil loading leads to a significant reduction in the unsteady lift and moment for the low reduced frequencies, but a significant decrease for the high reduced frequencies. Dissert. Abstr. N91-11671 Georgia Inst. of Tech., Atlanta. UNSTEADY VORTEX LATTICE AERODYNAMICS FOR ROTOR AEROELASTICITY IN HOVER AND IN FORWARD FLIGHT Ph.D. Thesis Kyung Min Yoo 1990 223 p Avail: Univ. Microfilms Order No. DA9023837 An unsteady vortex lattice method was applied and was shown to be an improvement over the classical two-dimensional quasi-steady aerodynamics. A number of code validation studies were carried out including comparison of the present method with experimental data and other methods. To develop a computational model of the rotor and its wake, a thin lifting surface and wake were discretized into bound, shed, and trailing vortex filaments fixed in a prescribed wake geometry. Also, to consider the unsteady wake aeromechanism among the blades, each blade's motion and deformation were treated. The unsteady induced inflow was calculated and compared with experiment and other inflow models. The results for a two-bladed hovering rotor clearly show the effects of the three-dimensional tip-relief effect and the unsteady wake dynamics effect of the near and returning wake undergoing various unsteady motions. The present thin lifting surface theory was also applied to a coupled flap-lag-torsion stability analysis for the hovering flight condition. The perturbed time histories of coupled flap-lag-torsional motions were analyzed by Fourier analyses to predict the damping and frequencies of particular modes. The overprediction of lead-lag damping by the two-dimensional quasi-steady aerodynamics is shown to be due to a lack of both three-dimensional tip-relief effects and unsteady wake dynamics of the near and returning wake. In forward flight, the present method is applied to aeroelastic response predictions for both low and high advance ratios. In low speed flight, there exists a strong blade-wake interaction which is diminished in high speed flight. Dissert. Abstr. N91-11672 Purdue Univ., West Lafayette, IN. Avail: Univ. Microfilms Order No. DA9018856 Three dimensional, unsteady, subsonic, and transonic flow through a single rotation propeller is studied. The unsteady loads on the blades are obtained by solving the full potential equations using an implicit time marching scheme. The purpose of the code is to provide a capability of doing propfan aeroelastic analysis in the nonlinear transonic regime. Results are shown for steady state aerodynamic loading, unsteady aerodynamic response to forced aeroelastic deformations, and free aeroelastic response. The aerodynamic analysis is based on a finite volume discretization of the potential equations. The scheme is fully implicit, with the resulting nonlinear algebraic equations being solved in conservation form by an approximately factored quasi-Newton iteration at each time step. The blade dynamics are based on the in vacuum modes and frequencies. The scheme uses a moving grid that conforms instantaneously to the deforming blade shape. Grids are generated, a priori, for the undeformed blade and for unit deformations in each of the in vacuum modes. Dynamic grids are then set by linear superposition based on the current state vector of the blade. The aeroelastic analysis of propfans was addressed in two ways: frequency domain analysis and time domain simulation. The frequency domain solution uses the generalized forces that are obtained from transfer function analysis. The results are compared directly to a linear panel method in terms of damping coefficients. Dissert. Abstr. N91-11673*# National Aeronautics and Space Administration. Langley Research Center, Hampton, VA. DEVELOPMENT OF UNSTRUCTURED GRID METHODS FOR STEADY AND UNSTEADY AERODYNAMIC ANALYSIS John T. Batina Nov. 1990 13 P Presented at the 17th Congress of the International Council of the Aeronautical Sciences, Stockholm, Sweden, 10-13 Sep. 1990 (NASA-TM-102730; NAS 1.15:102730; ICAS-90-6.9.4) Avail: NTIS HC/MF A03 CSCL 01A The current status of the development of unstructured grid methods in the Unsteady Aerodynamics Branch at NASA-Langley is described. These methods are being developed for steady and unsteady aerodynamic applications. The flow solvers that were developed for the solution of the unsteady Euler and Navier-Stokes equations are highlighted and selected results are given which demonstrate various features of the capability. The results demonstrate 2-D and 3-D applications for both steady and unsteady flows. Comparisons are also made with solutions obtained using a structured grid code and with experimental data to determine the accuracy of the unstructured grid methodology. These comparisons show good agreement which thus verifies the accuracy. Author N91-11674*# National Aeronautics and Space Administration. Presented at the 21st AIAA Fluid Dynamics, Plasma Dynamics and Lasers Conference, Seattle, WA, 18-20 Jun. 1990 Previously announced in IAA as A90-38777 (NASA-TM-102731; NAS 1.15:102731; AIAA-90-1649) Avail: NTIS HC/MF A02 CSCL 01A Improved algorithms for the solution of the 3-D time dependent Euler equations are presented for aerodynamic analysis involving unstructured dynamic meshes. The improvements were developed recently to the spatial and temporal discretizations used by unstructured grid flow solvers. The spatial discretization involves a flux split approach which is naturally dissipative and captures shock waves sharply with at most one grid point within the shock structure. The temporal discretization involves either an explicit time integration scheme using a multistage Runge-Kutta procedure or an implicit time integration scheme using a Gauss-Seidel relaxation procedure which is computationally efficient for either steady or unsteady flow problems. With the implicit Gauss-Seidel procedure, very large time steps may be used for rapid convergence to steady state, and the step size for unsteady cases may be selected for temporal accuracy rather than for numerical stability. Steady flow results are presented for both the NACA 0012 airfoil and the ONERA M6 wing to demonstrate applications of the new Euler solvers. A description of the Euler solvers is presented along with results and comparisons which assess the capability. Author K. B. M. Q. Zaman Oct. 1990 14 P (NASA-TM-103183; E-5563; NAS 1.15:103183; AIAA-90-4010) The effect of acoustic excitation on post-stalled flows over an airfoil, i.e., flows that are fully separated from near the leading edge, is investigated. The excitation results in a tendency towards reattachment, which is accompanied by an increased lift and reduced drag, although the flow may still remain fully separated. It is found that with increasing excitation amplitude, the effect becomes more pronounced but shifts to a Strouhal number which is much lower than that expected from linear, inviscid instability of the separated shear layer. Author N91-11676*# National Aeronautics and Space Administration. Lewis Research Center, Cleveland, OH. NUMERICAL STUDY OF UNSTEADY SHOCKWAVE REFLECTIONS USING AN UPWIND TVD SCHEME Andrew T. Hsu (Sverdrup Technology, Inc., Brook Park, OH.) and Meng-Sing Liou Aug. 1990 16 p (NASA-TM-103251; E-5680; NAS 1.15:103251) Avail: NTIS HC/MF A03 CSCL 01A An unsteady TVD Navier-Stokes solver was developed and applied to the problem of shock reflection on a circular cylinder. The obtained numerical results were compared with the Schlieren photos from an experimental study. These results show that the present computer code has the ability of capturing moving shocks. Author The use is discussed of gas injection from the surface of a hypervelocity vehicle such that the effective shape of the body is changed to obtain desired aerodynamic forces and thus desired vehicle control; this type of aerodynamic control represents a gas injection scheme which might possibly be used for control both inside and outside the atmosphere. Optimal trajectories and control were determined for trajectories and control multi-stage surface launched interceptors with ranges of several thousand miles and flight times of a few minutes; both propulsive and aerodynamic controls were considered, and trajectory optimization was based on minimizing the total interceptor mass ratio. In the aerodynamic control work, studies were made of gas injection into the boundary layer on one side of a thin wedge in steady hypersonic flow. To ascertain the magnitudes of the largest force changes available with blowing, gas injection rates large enough to cause the boundary layer to be blown off the body were considered. Conditions for the desired flight envelope are such that laminar flows could be considered. A fundamental conclusion reached is that relatively large aerodynamic force changes can be obtained with relatively small rates rates of gas injection. GRA A SUMMARY OF TRANSONIC NATURAL LAMINAR FLOW This report contains an analysis of the experimental results obtained from four supercritical natural laminar flow airfoils investigated in the NAE High Reynolds Number Test Facility. The airfoils have maximum thickness to chord ratios of 0.10, 0.13, 0.16 and 0.21 and were designed for a lift coefficient 0.6. Their design Mach numbers were 0.8, 0.76, 0.72 and 0.68 respectively and the design chord Reynolds number was 12.5 million. It was found that all the airfoils showed the presence of a drag bucket close to design conditions and long lengths (in some cases about 70 percent) of natural laminar flow at Reynolds number 6.7 million. The minimum drag for the airfoils was found to range from 0.0045 to 0.0051, representing far lower levels than any airfoil dominated by turbulent boundary layer. It is also indicated that, with transition fixed at about 10 percent chord, the drag levels were similar to other airfoils with turbulent boundary layers. GRA N91-11680# European Space Agency, Paris (France). Wolfgang Lorenz-Meyer (Deutsche Forschungs- und Versuchsanstalt fuer Luft- und Raumfahrt, Goettingen, Germany, F.R.) Aug. 1990 43 p Transl. into ENGLISH of Halbmodellmessungen an einem ONERA-Eichmodell im Transsonischen Windkanal Goettingen (Goettingen, Fed. Republic of Germany, DFLR), May 1989 44p Original language document was announced as N90-18370 (ESA-TT-1195; DLR-Mitt-89-20; ETN-90-98009) Avail: NTIS Force and pressure distribution measurements were carried out at three wing sections on a calibration model in a 1m by 1m transonic wind tunnel. The model was mounted on the half model balance without a splitter plate but with a 5 mm thick boundary layer trap. The results are compared with complete model tests. The coefficients cannot be adequately corrected by the use of linear correction rules. Some oil film photographs were prepared for flow visualization purposes. ESA N91-11681# Naval Postgraduate School, Monterey, CA. MAPPING THE AIRWAKE OF A MODEL DD-963 ALONG SPECIFIC HELICOPTER FLIGHT PATHS M.S. Thesis Gustav A. Anderson Dec. 1989 93 P (AD-A225327) Avail: NTIS HC/MF A05 CSCL 01/1 A continuation was made of flow visualization studies done in the NPS low speed environmental wind tunnel. The long term goal is to map the airwake around a ship model and scale to full size for the purpose of determining safe operating envelopes on non-aviation ships. This project used hot wire and hot film anemometry to establish a data base for helicopter approach paths at O deg, 30 deg port, and 30 deg starboard ship yaw angles. Calibration of the wind tunnel revealed that some turbulence generators, used in two previous studies, created excessive turbulence intensity levels and were subsequently removed. Analysis along the flight paths was done with and without the model in place. The comparison showed that turbulence intensity levels of up to 50 percent were experienced in the proximity of the flight deck. These levels fell by 40 to 50 percent within 1/4 ship length along all approach paths. The starboard yaw approach path contained the greatest turbulence magnitudes and the 0 deg yaw contained the least. GRA N91-11682*# National Aeronautics and Space Administration. AIRBORNE WIND SHEAR DETECTION AND WARNING Amos A. Spady, Jr., comp., Roland L. Bowles, comp., and Herbert (NASA-CP-10050-Pt-1; NAS 1.55:10050-Pt-1) Avail: NTIS HC/MF A15 CSCL 01C The Second Combined Manufacturers' and Technologists' Conference hosted jointly by NASA Langley (LaRC) and the Federal Aviation Administration (FAA) was held in Williamsburg, Virginia, on October 18 to 20, 1988. The purpose of the meeting was to transfer significant, ongoing results gained during the second year of the joint NASA/FAA Airborne Wind Shear Program to the technical industry and to pose problems of current concern to the combined group. It also provided a forum for manufacturers to review forward-look technology concepts and for technologists to gain an understanding of the problems encountered by the manufacturers during the development of airborne equipment and the FAA certification requirements. For individual titles, see N91-11683 through N91-11694. N91-11683*# American Airlines, Inc., Fort Worth, TX. Wallace M. Gillman In NASA, Langley Research Center, Airborne A brief review is given of daily operations in the airline business, with emphasis on the decisions made by pilots and the information used to make those decisions, Various wind shears are discussed as they affect daily operations. The discussion of tools focuses on airborne reactive and predictive systems. The escape maneuver used to fly out of a severe windshear is from a pilot's point of view. Author N91-11684*# Honeywell, Inc., Phoenix, AZ. Sperry Commercial FLIGHT EXPERIENCE WITH WINDSHEAR DETECTION Windshear alerts resulting from the Honeywell Windshear Detection and Guidance System are presented based on data from approximately 248,000 revenue flights at Piedmont Airlines. The data indicate that the detection system provides a significant benefit to the flight crew of the aircraft. In addition, nuisance and false alerts were found to occur at an acceptably low rate to maintain flight crew confidence in the system. Data from a digital flight recorder is also presented which shows the maximum and minimum windshear magnitudes recorded for a representative number of flights in February, 1987. The effect of the boundary Author layer of a steady state wind is also discussed. N91-11685*# Honeywell, Inc., Phoenix, AZ. Sperry Commercial Flight Systems Group. INTERFACE STANDARDS FOR INTEGRATED FORWARD-LOOKING/PREDICTIVE/REACTIVE WINDSHEAR SYSTEMS Mark M. McGlinchey In NASA, Langley Research Center, Airborne Forward-looking windshear systems are developing to a point (particularly the infrared sensors) where their interface with the cockpit and reactive windshear systems needs to be defined. As airlines retrofit their aircraft with reactive windshear systems, it is important to recognize that onboard windshear systems of the future will be a combination of both forward-looking and reactive elements. Today's reactive systems need to be built with the capability to interface with the forward-looking systems of tomorrow. This presentation is a first step at looking at the requirements and defining interface standards for integrated forward-looking and reactive windshear systems. Undoubtedly the requirements for interfacing these types of windshear systems will change as the technology changes. Author N91-11686*# National Aeronautics and Space Administration. Langley Research Center, Hampton, VA. HEAVY RAIN EFFECTS ON AIRPLANE PERFORMANCE R. E. Dunham, Jr., G. M. Bezos, B. A. Campbell, W. D. Mace, Jr., and W. E. Melson, Jr. (National Aeronautics and Space Administration. Wallops Flight Facility, Wallops Island, VA.) In its Airborne Wind Shear Detection and Warning Systems. Second Combined Manufactuters' and Technologists' Conference, Part 1 Jul. 1990 p 85-101 (For primary document see N91-11682 03-03) Avail: NTIS HC/MF A15 CSCL 01C The objective is to determine if the aerodynamic characteristics of an airplane are altered while flying in the rain. Wind-tunnel tests conducted at the NASA Langley Research Center (LaRC) have shown losses in maximum lift, reduction in stall angle, and increases in drag when a wing is placed in a simulated rain spray. For these tests the water spray concentration used represented a very heavy rainfall. A lack of definition of the scaling laws for aerodynamic testing in a two-phase, two-component flow makes interpolation of the wind-tunnel test uncertain. Tests of a large-scale wing are to be conducted at the LaRC. The large-scale wing is mounted on top of the Aircraft Landing Dynamics Facility (ALDF) carriage. This carriage (which is 70-foot long, 30-foot wide, and 30-foot high) is propelled with the wing model attached down a 3000-foot long test track by a water jet at speeds of up to 170 |