Engineering topics

Scallion, William I.

Publications and source records attributed to Scallion, William I..

14 records

NASA NTRS2009

Aerodynamic Interference Due to MSL Reaction Control System

An investigation of effectiveness of the reaction control system (RCS) of Mars Science Laboratory (MSL) entry capsule during atmospheric flight has been conducted. The reason for the investigation is that MSL is designed to fly a lifting actively guided entry with hypersonic bank maneuvers, therefore an understanding of RCS effectiveness is required. In the course of the study several jet configurations were evaluated using Langley Aerothermal Upwind Relaxation Algorithm (LAURA) code, Data Parallel Line Relaxation (DPLR) code, Fully Unstructured 3D (FUN3D) code and an Overset Grid Flowsolver (OVERFLOW) code. Computations indicated that some of the proposed configurations might induce aero-RCS interactions, sufficient to impede and even overwhelm the intended control torques. It was found that the maximum potential for aero-RCS interference exists around peak dynamic pressure along the trajectory. Present analysis largely relies on computational methods. Ground testing, flight data and computational analyses are required to fully understand the problem. At the time of this writing some experimental work spanning range of Mach number 2.5 through 4.5 has been completed and used to establish preliminary levels of confidence for computations. As a result of the present work a final RCS configuration has been designed such as to minimize aero-interference effects and it is a design baseline for MSL entry capsule.

Dyakonov, Artem A.↗

NASA NTRS2007

Simulation and Analyses of Stage Separation of Two-Stage Reusable Launch Vehicles

NASA has initiated the development of methodologies, techniques and tools needed for analysis and simulation of stage separation of next generation reusable launch vehicles. As a part of this activity, ConSep simulation tool is being developed which is a MATLAB-based front-and-back-end to the commercially available ADAMS(Registerd TradeMark) solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the application of ConSep to the simulation and analysis of staging maneuvers of two-stage-to-orbit (TSTO) Bimese reusable launch vehicles, one staging at Mach 3 and the other at Mach 6. The proximity and isolated aerodynamic database were assembled using the data from wind tunnel tests conducted at NASA Langley Research Center. The effects of parametric variations in mass, inertia, flight path angle, altitude from their nominal values at staging were evaluated. Monte Carlo runs were performed for Mach 3 staging to evaluate the sensitivity to uncertainties in aerodynamic coefficients.

Pamadi, Bandu N.↗

NASA NTRS2005

Simulation and Analyses of Stage Separation Two-Stage Reusable Launch Vehicles

NASA has initiated the development of methodologies, techniques and tools needed for analysis and simulation of stage separation of next generation reusable launch vehicles. As a part of this activity, ConSep simulation tool is being developed which is a MATLAB-based front-and-back-end to the commercially available ADAMS(registered Trademark) solver, an industry standard package for solving multi-body dynamic problems. This paper discusses the application of ConSep to the simulation and analysis of staging maneuvers of two-stage-to-orbit (TSTO) Bimese reusable launch vehicles, one staging at Mach 3 and the other at Mach 6. The proximity and isolated aerodynamic database were assembled using the data from wind tunnel tests conducted at NASA Langley Research Center. The effects of parametric variations in mass, inertia, flight path angle, altitude from their nominal values at staging were evaluated. Monte Carlo runs were performed for Mach 3 staging to evaluate the sensitivity to uncertainties in aerodynamic coefficients.

Pamadi, Bandu N.↗

NASA NTRS2005

Experimental Stage Separation Tool Development in NASA Langley's Aerothermodynamics Laboratory

As part of the research effort at NASA in support of the stage separation and ascent aerothermodynamics research program, proximity testing of a generic bimese wing-body configuration was conducted in NASA Langley's Aerothermodynamics Laboratory in the 20-Inch Mach 6 Air Tunnel. The objective of this work is the development of experimental tools and testing methodologies to apply to hypersonic stage separation problems for future multi-stage launch vehicle systems. Aerodynamic force and moment proximity data were generated at a nominal Mach number of 6 over a small range of angles of attack. The generic bimese configuration was tested in a belly-to-belly and back-to-belly orientation at 86 relative proximity locations. Over 800 aerodynamic proximity data points were taken to serve as a database for code validation. Longitudinal aerodynamic data generated in this test program show very good agreement with viscous computational predictions. Thus a framework has been established to study separation problems in the hypersonic regime using coordinated experimental and computational tools.

Murphy, Kelly J.↗

NASA NTRS2004

Overview of Transonic to Hypersonic Stage Separation Tool Development for Multi-Stage-to-Orbit Concepts

An overview of research efforts at NASA in support of the stage separation and ascent aerothermodynamics research program is presented. The objective of this work is to develop a synergistic suite of experimental, computational, and engineering tools and methods to apply to vehicle separation across the transonic to hypersonic speed regimes. Proximity testing of a generic bimese wing-body configuration is on-going in the transonic (Mach numbers 0.6, 1.05, and 1.1), supersonic (Mach numbers 2.3, 3.0, and 4.5) and hypersonic (Mach numbers 6 and 10) speed regimes in four wind tunnel facilities at the NASA Langley Research Center. An overset grid, Navier-Stokes flow solver has been enhanced and demonstrated on a matrix of proximity cases and on a dynamic separation simulation of the bimese configuration. Steady-state predictions with this solver were in excellent agreement with wind tunnel data at Mach 3 as were predictions via a Cartesian-grid Euler solver. Experimental and computational data have been used to evaluate multi-body enhancements to the widely-used Aerodynamic Preliminary Analysis System, an engineering methodology, and to develop a new software package, SepSim, for the simulation and visualization of vehicle motions in a stage separation scenario. Web-based software will be used for archiving information generated from this research program into a database accessible to the user community. Thus, a framework has been established to study stage separation problems using coordinated experimental, computational, and engineering tools.

Murphy, Kelly J.↗

NASA NTRS2004

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Brauckmann, Gregory J.↗

NASA NTRS2004

CFD Approaches for Simulation of Wing-Body Stage Separation

A collection of computational fluid dynamics tools and techniques are being developed and tested for application to stage separation and abort simulation for next-generation launch vehicles. In this work, an overset grid Navier-Stokes flow solver has been enhanced and demonstrated on a matrix of proximity cases and on a dynamic separation simulation of a belly-to-belly wing-body configuration. Steady cases show excellent agreement between Navier-Stokes results, Cartesian grid Euler solutions, and wind tunnel data at Mach 3. Good agreement has been obtained between Navier-Stokes, Euler, and wind tunnel results at Mach 6. An analysis of a dynamic separation at Mach 3 demonstrates that unsteady aerodynamic effects are not important for this scenario. Results provide an illustration of the relative applicability of Euler and Navier-Stokes methods to these types of problems.

Buning, Pieter G.↗

NASA NTRS2003

Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

Brauckman, Gregory J.↗

NASA NTRS1999

Aerodynamic Characteristics and Control Effectiveness of the HL-20 Lifting Body Configuration at Mach 10 in Air

A 0.0196-scale model of the HL-20 lifting-body, one of several configurations proposed for future crewed spacecraft, was tested in the Langley 31-Inch Mach 10 Tunnel. The purpose of the tests was to determine the effectiveness of fin-mounted elevons, a lower surface flush-mounted body flap, and a flush-mounted yaw controller at hypersonic speeds. The nominal angle-of-attack range, representative of hypersonic entry, was 2 deg to 41 deg, the sideslip angles were 0 deg, 2 deg, and -2 deg, and the test Reynolds number was 1.06 x 10 E6 based on model reference length. The aerodynamic, longitudinal, and lateral control effectiveness along with surface oil flow visualizations are presented and discussed. The configuration was longitudinally and laterally stable at the nominal center of gravity. The primary longitudinal control, the fin-mounted elevons, could not trim the model to the desired entry angle of attack of 30 deg. The lower surface body flaps were effective for roll control and the associated adverse yawing moment was eliminated by skewing the body flap hinge lines. A yaw controller, flush-mounted on the lower surface, was also effective, and the associated small rolling moment was favorable.

Scallion, William I.↗

NASA NTRS1991

Effects of nozzle exit geometry and pressure ratio on plume shape for nozzles exhausting into quiescent air

The effects of varying the exit geometry on the plume shapes of supersonic nozzles exhausting into quiescent air at several exit-to-ambient pressure ratios are given. Four nozzles having circular throat sections and circular, elliptical and oval exit cross sections were tested and the exit plume shapes are compared at the same exit-to-ambient pressure ratios. The resulting mass flows were calculated and are also presented.

Scallion, William I.↗

NASA NTRS1989

Langley proposed advanced hypervelocity aerophysics facility - A status report

A ground-based facility capable of performing flight tests on relatively large highly instrumented models and scaled vehicle components at velocities and densities representative of a hypervelocity flight in earth and planetary atmospheres is reviewed. This facility proposed by the Langley Research Center is based on a launcher, a test chamber, and a model impact/deceleration chamber. It would initially utilize existing light-gas gun launcher technology scaled to 4 times present launcher size. It is planned to enhance its velocity and model size capability either by an electromagnetic launcher or a ram accelerator.

Witcofski, Robert D.↗

NASA NTRS1989

Advanced Hypervelocity Aerophysics Facility Workshop

The primary objective of the workshop was to obtain a critical assessment of a concept for a large, advanced hypervelocity ballistic range test facility powered by an electromagnetic launcher, which was proposed by the Langley Research Center. It was concluded that the subject large-scale facility was feasible and would provide the required ground-based capability for performing tests at entry flight conditions (velocity and density) on large, complex, instrumented models. It was also concluded that advances in remote measurement techniques and particularly onboard model instrumentation, light-weight model construction techniques, and model electromagnetic launcher (EML) systems must be made before any commitment for the construction of such a facility can be made.

Witcofski, Robert D.↗

NASA NTRS1963

Full-Scale Wind-Tunnel Investigation of the Drag Characteristics of an HU2K Helicopter Fuselage

An investigation was conducted in the Langley full-scale tunnel to determine the drag characteristics of the HU2K helicopter fuselage. The effects of body shape, engine operation, appendages, and leakage on the model drag were determined. The results of the tests showed that the largest single contribution to the parasite drag was that of the rotor hub installation which produced about 80 percent of the drag of the sealed and faired production body. Fairings on the rotor hub and blade retentions, or a cleaned-up hub and retentions, appeared to be the most effective single modifications tested. The total drag of all protuberances and air leakage also contributed a major part of the drag - an 83-percent increase over the drag of the sealed and faired production body. An additional increment of drag was caused by the basic shape of the fuselage - 19 percent more than the drag obtained when the fuselage shape was extensively refaired. Another sizable increment of drag was caused by the engine oil-cooler exit which gave a drag of 8 percent of that of the sealed and faired production body.

Scallion, William I.↗

NASA NTRS1950

Data from Tests of a 1/5-Scale Model of a Proposed High-Speed Submarine in the Langley Full-Scale Tunnel

Tests of a 1/5 scale model of a proposed 153-foot high-speed submarine have been conducted in the Langley full-scale tunnel at the request of the Bureau of Ships, Department of the Navy. The test program included: (1) force tests to determine the drag, control effectiveness, and static stability characteristics for a number of model configurations, both in pitch and in yaw, (2) pressure measurements to determine the boundary-layer conditions and flow characteristics in the region of the propeller, and (3) an investigation of the effects of propeller operation on the model aerodynamic characteristics. In response to oral requests from the Bureau of Ships representatives t hat the basic data obtained in these tests be made available to them as rapidly as possible, this data report has been prepared to present some of the more pertinent results. All test results given in the present paper are for the propeller-removed condition and were obtained at a Reynolds number of approximately 22,300,000 based on model length.

Cocke, Bennie W.↗