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Bobbitt, Percy J.

Publications and source records attributed to Bobbitt, Percy J..

18 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

Bobbitt, Percy J.↗
NASA NTRS2011

Wedge and Conical Probes for the Instantaneous Measurement of Free-Stream Flow Quantities at Supersonic Speeds

Wedge and conical shaped probes for the measurement of free-stream flow quantities at supersonic speeds have been tested in both wind tunnel and flight. These probes have improved capabilities over similar ones used in the past. Through the use of miniature pressure sensors, that are located inside the probes, they are able to provide instantaneous measurements of a time-varying environment. Detailed herein are the results of the tests in NASA Langley Researcher Center s Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8 and 2.0, as well as flight tests carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15 aircraft up to Mach numbers of 1.9. In the flight tests the probes were attached to a fixture on the underside of the F-15 fuselage. Problems controlling the velocity of the flow through the conical probe, required for accurate temperature measurements, are noted, as well as some calibration problems of the miniature pressure sensors that impact the accuracy of the measurements.

Bobbitt, Percy J.↗
NASA NTRS2011

Focused and Steady-State Characteristics of Shaped Sonic Boom Signatures: Prediction and Analysis

The objective of this study is to examine the effect of flight, at off-design conditions, on the propagated sonic boom pressure signatures of a small "low-boom" supersonic aircraft. The amplification, or focusing, of the low magnitude "shaped" signatures produced by maneuvers such as the accelerations from transonic to supersonic speeds, climbs, turns, pull-up and pushovers is the concern. To analyze these effects, new and/or improved theoretical tools have been developed, in addition to the use of existing methodology. Several shaped signatures are considered in the application of these tools to the study of selected maneuvers and off-design conditions. The results of these applications are reported in this paper as well as the details of the new analytical tools. Finally, the magnitude of the focused boom problem for "low boom" supersonic aircraft designs has been more accurately quantified and potential "mitigations" suggested. In general, "shaped boom" signatures, designed for cruise flight, such as asymmetric and symmetric flat-top and initial-shock ramp waveforms retain their basic shape during transition flight. Complex and asymmetric and symmetric initial shock ramp waveforms provide lower magnitude focus boom levels than N-waves or asymmetric and symmetric flat-top signatures.

Maglieri, Domenic J.↗
NASA NTRS2011

Equivalent Longitudinal Area Distributions of the B-58 and XB-70-1 Airplanes for Use in Wave Drag and Sonic Boom Calculations

A detailed geometric description, in wave drag format, has been developed for the Convair B-58 and North American XB-70-1 delta wing airplanes. These descriptions have been placed on electronic files, the contents of which are described in this paper They are intended for use in wave drag and sonic boom calculations. Included in the electronic file and in the present paper are photographs and 3-view drawings of the two airplanes, tabulated geometric descriptions of each vehicle and its components, and comparisons of the electronic file outputs with existing data. The comparisons include a pictorial of the two airplanes based on the present geometric descriptions, and cross-sectional area distributions for both the normal Mach cuts and oblique Mach cuts above and below the vehicles. Good correlation exists between the area distributions generated in the late 1950s and 1960s and the present files. The availability of these electronic files facilitates further validation of sonic boom prediction codes through the use of two existing data bases on these airplanes, which were acquired in the 1960s and have not been fully exploited.

Tinetti, Ana F.↗
NASA NTRS2002

Computational Predictions of the Performance Wright 'Bent End' Propellers

Computational analysis of two 1911 Wright brothers 'Bent End' wooden propeller reproductions have been performed and compared with experimental test results from the Langley Full Scale Wind Tunnel. The purpose of the analysis was to check the consistency of the experimental results and to validate the reliability of the tests. This report is one part of the project on the propeller performance research of the Wright 'Bent End' propellers, intend to document the Wright brothers' pioneering propeller design contributions. Two computer codes were used in the computational predictions. The FLO-MG Navier-Stokes code is a CFD (Computational Fluid Dynamics) code based on the Navier-Stokes Equations. It is mainly used to compute the lift coefficient and the drag coefficient at specified angles of attack at different radii. Those calculated data are the intermediate results of the computation and a part of the necessary input for the Propeller Design Analysis Code (based on Adkins and Libeck method), which is a propeller design code used to compute the propeller thrust coefficient, the propeller power coefficient and the propeller propulsive efficiency.

Wang, Xiang-Yu↗
NASA NTRS1994

Experimental studies of transonic flow field near a longitudinally slotted wind tunnel wall

The results of detailed parametric experiments are presented for the near-wall flow field of a longitudinally slotted transonic wind tunnel. Existing data are reevaluated and new data obtained in the Langley 6- by 19-inch Transonic Wind Tunnel are presented and analyzed. In the experiments, researchers systematically investigate many pertinent wall-geometry variables such as the wall openness and the number of slots along with the free stream Mach number and model angle of attack. Flow field surveys on the plane passing through the centerline of the slot were conducted and are presented. The effects of viscosity on the slot flow are considered in the analysis. The present experiments, combined with those of previous investigations, give a more complete physical characterization of the flow near and through the slotted wall of a transonic wind tunnel.

Everhart, Joel L.↗
NASA NTRS1992

Application of computational fluid dynamics and laminar flow technology for improved performance and sonic boom reduction

A discussion is given of the many factors that affect sonic booms with particular emphasis on the application and development of improved computational fluid dynamics (CFD) codes. The benefits that accrue from interference (induced) lift, distributing lift using canard configurations, the use of wings with dihedral or anhedral and hybrid laminar flow control for drag reduction are detailed. The application of the most advanced codes to a wider variety of configurations along with improved ray-tracing codes to arrive at more accurate and, hopefully, lower sonic booms is advocated. Finally, it is speculated that when all of the latest technology is applied to the design of a supersonic transport it will be found environmentally acceptable.

Bobbitt, Percy J.↗
NASA NTRS1992

Overview of feasibility study on conducting overflight measurements of shaped sonic boom signatures using RPV's

Before beginning this presentation, it is appropriate to acknowledge the sincere interest and financial support provided by the NASA LaRC under contract NAS9-17900. An outline of the material to be used in the present paper is given. It begins with a indication of the origin and objectives of the feasibility study. This is followed by a discussion of various simulation methods of establishing the persistence of shaped sonic boom signatures to large distances including the use of recoverable RPV/drones. The desirable features to be sought out in an RPV along with a rationale for the selection of a 'shaped' sonic boom signature will be addressed. Three candidate vehicles are examined as to their suitability with respect to a number of factors, in particular, modifiability. Area distributions and associated sonic boom signatures of the basic and modified Firebee vehicle will also be shown. An indication of the scope of the proposed wind tunnel and flight test programs will be presented including measurement technologies and predicted waveforms. Finally, some remarks will be made summarizing the study and highlighting the key findings.

Maglieri, Domenic J.↗
NASA NTRS1992

Results for the hybrid laminar flow control experiment conducted in the NASA Langley 8-foot transonic pressure tunnel on a 7-foot chord model

A description is given of the development of, and results from, the hybrid laminar flow control (HLFC) experiment conducted in the NASA LaRC 8 ft Transonic Pressure Tunnel on a 7 ft chord, 23 deg swept model. The methods/codes used to obtain the contours of the HLFC model surface and to define the suction requirements are outlined followed by a discussion of the model construction, suction system, instrumentation, and some example results from the wind tunnel tests. Included in the latter are the effects of Mach number, suction level, and the extent of suction. An assessment is also given of the effect of the wind tunnel environment on the suction requirements. The data show that, at or near the design Mach number, large extents of laminar flow can be achieved with suction mass flows over the first 25 percent, or less, of the chord. Top surface drag coefficients with suction extending from the near leading edge to 20 percent of the chord were approximately 40 percent lower than those obtained with no suction. The results indicate that HLFC can be designed for transonic speeds with lift and drag coefficients approaching those of LFC designs but with much smaller extents and levels of suction.

Bobbitt, Percy J.↗
NASA NTRS1992

The Langley 8-ft transonic pressure tunnel laminar-flow-control experiment

An account is given of the considerations involved in selecting the NASA-Langley transonic pressure tunnel's design and test parameters, as well as its liner and a swept wing for laminar flow control (LFC) experimentation. Attention is given to the types and locations of the instrumentation employed. Both slotted and perforated upper surfaces were tested with partial- and full-chord suction; representative results are presented for all.

Bobbitt, Percy J.↗
NASA NTRS1990

Transition research opportunities at subsonic and transonic speeds

A number of opportunities available to experimentalists in the subsonic- and transonic-flow regime is discussed, including the needs for improving the quality of flow in wind tunnels and the techniques used to measure it. Attention is focused on uncertainties in transition data obtained in wind-tunnel experiments, methods of flow quality improvements, relationships between Reynolds-number effects and flow quality, spanwise pressure gradient, and active transition control. Advances in stability theory and Navier-Stokes simulations are outlined, along with measurement/instrument needs, advantages of momentum-thickness transition criteria, and the effects of laminar flow on the induced drag of a wing.

Bobbitt, Percy J.↗
NASA NTRS1989

Instrumentation advances for transonic testing

New and improved instrumentation, like new and improved wind tunnels, provide capabilities which stimulate innovative research and discovery. During the past few years there have been a number of instrumentation developments which have aided and abetted the acquisition of more accurate aerodynamic data and have led to new physical insights as well. Some of these advances are reviewed, particularly in the area of thin film gages, hot wire anemometry, and laser instrumentation. A description is given of the instruments and/or techniques and some sample results are shown.

Bobbitt, Percy J.↗
NASA NTRS1988

The pros and cons of code validation

Computational and wind tunnel error sources are examined and quantified using specific calculations of experimental data, and a substantial comparison of theoretical and experimental results, or a code validation, is discussed. Wind tunnel error sources considered include wall interference, sting effects, Reynolds number effects, flow quality and transition, and instrumentation such as strain gage balances, electronically scanned pressure systems, hot film gages, hot wire anemometers, and laser velocimeters. Computational error sources include math model equation sets, the solution algorithm, artificial viscosity/dissipation, boundary conditions, the uniqueness of solutions, grid resolution, turbulence modeling, and Reynolds number effects. It is concluded that, although improvements in theory are being made more quickly than in experiments, wind tunnel research has the advantage of the more realistic transition process of a right turbulence model in a free-transition test.

Bobbitt, Percy J.↗
NASA NTRS1988

The pros and cons of code validation

Computational and wind tunnel error sources are examined and quantified using specific calculations or experimental data, and a substantial comparison of theoretical and experimental results, or a code validation, is discussed. Wind tunnel error sources considered include wall interference, sting effects, Reynolds number effects, flow quality and transition, and instrumentation such as strain gage balances, electronically scanned pressure systems, hot film gages, hot wire anemometers, and laser velocimeters. Computational error sources include math model equation sets, the solution algorithm, artificial viscosity/dissipation, boundary conditions, the uniqueness of solutions, grid resolution, turbulence modeling, and Reynolds number effects. It is concluded that although improvements in theory are being made more quickly than in experiments, wind tunnel research has the advantage of the more realistic transition process of a right turbulence model in a free-transition test.

Bobbitt, Percy J.↗
NASA NTRS1987

A review of transonic code validation research activities

It is established that Computational Fluid Dynamics (CFD) code validation is an integral part of all flow testing. Specific attention is given to the development of new methods/instruments to obtain time varying 3-D data. Additionally, a discussion concerning wind tunnels and small test facilities is presented. All results and data are presented in viewgraph form.

Bobbitt, Percy J.↗
NASA NTRS1986

Computational fluid dynamics overview

The hierarchy of codes; boundary layers and turbulence models; uniqueness and artificial viscosity; efficiency of solution algorithms; data display and analysis; and cost of experimental validation are briefly examined.

Bobbitt, Percy J.↗
NASA NTRS1959

Tables for the Rapid Estimation of Downwash and Sidewash Behind Wings Performing Various Motions at Supersonic Speeds

Equations for the downwash and sidewash due to supersonic yawed and unswept horseshoe vortices have been utilized in formulating tables and charts to permit a rapid estimation of the flow velocities behind wings performing various steady motions. Tabulations are presented of the downwash and sidewash in the wing vertical plane of symmetry due to a unit-strength yawed horseshoe vortex located at 20 equally spaced spanwise positions along lifting lines of various sweeps. (The bound portion of the yawed vortex is coincident with the lifting line.) Charts are presented for the purpose of estimating the spanwise variations of the flow-field velocities and give longitudinal variations of the downwash and sidewash at a nuMber of vertical and spanwise locations due to a unit-strength unswept horseshoe vortex. Use of the tables and charts to calculate wing downwash or sidewash requires a knowledge of the wing spanwise distribution of circulation. Sample computations for the rolling sidewash and angle-of-attack downwash behind a typical swept wing are presented to demonstrate the use of the tables and charts.

Bobbitt, Percy J.↗