Paul Barnhart - Academia.edu (original) (raw)
Papers by Paul Barnhart
A numerical, and experimental investigation to study the effects of flow distortion on a Mass Flo... more A numerical, and experimental investigation to study the effects of flow distortion on a Mass Flow Plug (MFP) used to control and measure mass-flow during an inlet test has been conducted. The MFP was first calibrated using the WIND-US flow solver for uniform (undistorted) inflow conditions. These results are shown to compare favorably with an experimental calibration under similar conditions. The effects of distortion were investigated by imposing distorted flow conditions taken from an actual inlet test to the inflow plane of the numerical simulation. The computational fluid dynamic (CFD) based distortion study only showed the general trend in mass flow rate. The study used only total pressure as the upstream boundary condition, which was not enough to define the flow. A better simulation requires knowledge of the turbulence structure and a specific distortion pattern over a range of plug positions. It is recommended that future distortion studies utilize a rake with at least the ...
The design characteristics of an inlet very much depend on whether the inlet is to be flown at su... more The design characteristics of an inlet very much depend on whether the inlet is to be flown at subsonic, supersonic, or hypersonic speed. Whichever the case, the primary function of an inlet is to deliver free-stream air to the engine face at the highest stagnation pressure possible and with the lowest possible variation in both stagnation pressure and temperature. At high speeds, this is achieved by a system of oblique and/or normal shock waves, and possibly some isentropic compression. For both subsonic and supersonic flight, current design practice indicates that the inlet should deliver the air to the engine face at approximately Mach 0.45. As a result, even for flight in the high subsonic regime, the inlet must retard (or diffuse) the air substantially. Second, the design of an inlet is influenced largely by the compromise between high performance and low weight. This compromise involves tradeoffs between the mission requirements, flight trajectory, airframe aerodynamics, engin...
A series of analyses have been developed which permit the calculation of the performance of commo... more A series of analyses have been developed which permit the calculation of the performance of common inlet designs. The methods presented are useful for determining the inlet weight flows, total pressure recovery, and aerodynamic drag coefficients for given inlet geometric designs. Limited geometric input data is required to use this inlet performance prediction methodology. The analyses presented here may also be used to perform inlet preliminary design studies. The calculated inlet performance parameters may be used in subsequent engine cycle analyses or installed engine performance calculations for existing uninstalled engine data.
The implementation of a magnetic suspension system in the NASA Glenn Research Center (GRC) 225 cm... more The implementation of a magnetic suspension system in the NASA Glenn Research Center (GRC) 225 cm2 Supersonic Wind Tunnel would be a powerful test technique that could accurately determine the dynamic stability of blunt body entry vehicles with no sting interference. This paper explores initial design challenges to be evaluated before implementation, including defining the lowest possible operating dynamic pressure and corresponding model size, developing a compatible video analysis technique, and incorporating a retractable initial support sting.
As part of a contract with the NASA Lewis Research Center, a simple, accurate method of predictin... more As part of a contract with the NASA Lewis Research Center, a simple, accurate method of predicting the performance characteristics of a nozzle design has been developed for use in conceptual design studies. The Nozzle Performance Analysis Code (NPAC) can predict the on- and off-design performance of axisymmetric or two-dimensional convergent and convergent-divergent nozzle geometries. NPAC accounts for the effects of overexpansion or underexpansion, flow divergence, wall friction, heat transfer, and small mass addition or loss across surfaces when the nozzle gross thrust and gross thrust coefficient are being computed. NPAC can be used to predict the performance of a given nozzle design or to develop a preliminary nozzle system design for subsequent analysis. The input required by NPAC consists of a simple geometry definition of the nozzle surfaces, the location of key nozzle stations (entrance, throat, exit), and the nozzle entrance flow properties. NPAC performs three analysis &qu...
A series of experiments were performed to investigate the effects of Mach number variation on the... more A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from...
Methods The main aspect of this project is to determine the feasibility of creating a hyper spect... more Methods The main aspect of this project is to determine the feasibility of creating a hyper spectral instrument on a miniaturized scale. A Cubesat unit is 10cm x 10cm x 10cm. This is the goal size of the instrument while still gaining the desired spectral and spatial resolution. Several Cubesat units can be ajoined to allow for larger payload, but these compartments would just be used for attitude control and power supply.
A Mathematical theory is developed to perform the calculations necessary to determine the wave dr... more A Mathematical theory is developed to perform the calculations necessary to determine the wave drag for slender bodies of non-circular cross section. The derivations presented in this report are based on extensions to supersonic linearized small perturbation theory. A numerical scheme is presented utilizing Fourier decomposition to compute the pressure coefficient on and about a slender body of arbitrary cross section.
24th Joint Propulsion Conference, 1988
For Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic thrQugh-fl... more For Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic thrQugh-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one-dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drags and size and weight. NOMENCLATURE A area BPR bypass ratio C D drag coefficient M Mach number
A series of experiments were performed to investigate the effects of Mach number variation on the... more A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from the dynamic pressure transducers were found to collapse to self-similar distributions as a function of the distance perpendicular to the separation line. The standard deviation of the pressure signals showed initial peaked distribution, with the maximum standard deviation point corresponding to the location of the separation line at Mach number 3.0 to 5.0. At Mach 2.0 the maximum standard deviation point was found to occur significantly upstream of the separation line. The intermittency distributions of the separation shock wave motion were found to be self-similar profiles for all Mach numbers. The intermittent region length was found to increase with Mach number and decrease with interaction sweepback angle. For Mach numbers 3.0 to 5.0 the separation line was found to correspond to high intermittencies or equivalently to the downstream locus of the separation shock wave motion. The Mach 2.0 tests, however, showed that the intermittent region occurs significantly upstream of the separation line. Power spectral densities measured in the intermittent regions were found to have self-similar frequency distributions when compared as functions of a Strouhal number for all Mach numbers and interaction sweepback angles. The maximum zero-crossing frequencies were found to correspond with the peak frequencies in the power spectra measured in the intermittent region.
A simple and accurate nozzle performance analysis methodology has been developed. The geometry mo... more A simple and accurate nozzle performance analysis methodology has been developed. The geometry modeling requirements are minimal and very flexible, thus allowing rapid design evaluations. The solution techniques accurately couple: continuity, momentum, energy, state, and other relations which permit fast and accurate calculations of nozzle gross thrust. The control volume and internal flow analyses are capable of accounting for the effects of: over/under expansion, flow divergence, wall friction, heat transfer, and mass addition/loss across surfaces. The results from the nozzle performance methodology are shown to be in excellent agreement with experimental data for a variety of nozzle designs over a range of operating conditions. preliminary nozzle system design, and subsequent performance analyses. Nozzle performance is typically described by determining three quantities: accepted engine airflow, WT, gross thrust coefficient, Cm, and aerodynamic drag coefficient, Co. It is also very important to be able to characterize nozzle performance over the entire vehicle flight and engine operation range, not just at the nozzle design point. The methodology presented
51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015
51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015
Aircraft Design and Operations Meeting, 1989
A study is undertaken to investigate the engine airframe integration effects for supersonic throu... more A study is undertaken to investigate the engine airframe integration effects for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic throughflow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic throughflow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic throughflow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum liftto-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from i.i0 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration. NOMENCLATURE BPR bypass ratio CD drag coefficient, D/gS CL lift coefficient, L/qS D drag L lift M Mach number q dynamic pressure S reference area Subscrip__ f friction w wave INTRODUCTION In perusing the technologies required for efficient long range supersonic cruise aircraft, NASA has sponsored a number of studies to identify suitable propulsion system concepts. In the past, conventional and variable cycle engines were considered the most likely candidates. With renewed *Supervisor, Aerospace Analysis Section, Member AIAA Another desirable aspect of the engine cycle is that the bypass ratio decreases with increasing flight Math number, providing higher cruise thrust. Reference 3 shows that much of the improved specific fuel consumption for supersonic through-flow fan engines results from potential improvements in installation efficiencies for long range supersonic cruise applications. The possible improvements in inlet performance with supersonic through-flow fan engines has been shown in reference 4. This study addresses the engine airframe integration characteristics for supersonic through-flow fan propulsion systems, six different supersonic through-flow fan engine installations are examined in this study. The effects of engine size, nacelle contouring, nacelle placement, and approximate bypass plume modeling are investigated. Additionally, the supersonic through-flow fan (STFF) nacelle installations are compared with a conventional turbine bypass engine (TBE) nacelle installation on the basis of installed wave drags and maximum lift to drag ratios for complete engine airframe the
51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2013
A numerical, and experimental investigation to study the effects of flow distortion on a Mass Flo... more A numerical, and experimental investigation to study the effects of flow distortion on a Mass Flow Plug (MFP) used to control and measure mass-flow during an inlet test has been conducted. The MFP was first calibrated using the WIND-US flow solver for uniform (undistorted) inflow conditions. These results are shown to compare favorably with an experimental calibration under similar conditions. The effects of distortion were investigated by imposing distorted flow conditions taken from an actual inlet test to the inflow plane of the numerical simulation. The computational fluid dynamic (CFD) based distortion study only showed the general trend in mass flow rate. The study used only total pressure as the upstream boundary condition, which was not enough to define the flow. A better simulation requires knowledge of the turbulence structure and a specific distortion pattern over a range of plug positions. It is recommended that future distortion studies utilize a rake with at least the ...
The design characteristics of an inlet very much depend on whether the inlet is to be flown at su... more The design characteristics of an inlet very much depend on whether the inlet is to be flown at subsonic, supersonic, or hypersonic speed. Whichever the case, the primary function of an inlet is to deliver free-stream air to the engine face at the highest stagnation pressure possible and with the lowest possible variation in both stagnation pressure and temperature. At high speeds, this is achieved by a system of oblique and/or normal shock waves, and possibly some isentropic compression. For both subsonic and supersonic flight, current design practice indicates that the inlet should deliver the air to the engine face at approximately Mach 0.45. As a result, even for flight in the high subsonic regime, the inlet must retard (or diffuse) the air substantially. Second, the design of an inlet is influenced largely by the compromise between high performance and low weight. This compromise involves tradeoffs between the mission requirements, flight trajectory, airframe aerodynamics, engin...
A series of analyses have been developed which permit the calculation of the performance of commo... more A series of analyses have been developed which permit the calculation of the performance of common inlet designs. The methods presented are useful for determining the inlet weight flows, total pressure recovery, and aerodynamic drag coefficients for given inlet geometric designs. Limited geometric input data is required to use this inlet performance prediction methodology. The analyses presented here may also be used to perform inlet preliminary design studies. The calculated inlet performance parameters may be used in subsequent engine cycle analyses or installed engine performance calculations for existing uninstalled engine data.
The implementation of a magnetic suspension system in the NASA Glenn Research Center (GRC) 225 cm... more The implementation of a magnetic suspension system in the NASA Glenn Research Center (GRC) 225 cm2 Supersonic Wind Tunnel would be a powerful test technique that could accurately determine the dynamic stability of blunt body entry vehicles with no sting interference. This paper explores initial design challenges to be evaluated before implementation, including defining the lowest possible operating dynamic pressure and corresponding model size, developing a compatible video analysis technique, and incorporating a retractable initial support sting.
As part of a contract with the NASA Lewis Research Center, a simple, accurate method of predictin... more As part of a contract with the NASA Lewis Research Center, a simple, accurate method of predicting the performance characteristics of a nozzle design has been developed for use in conceptual design studies. The Nozzle Performance Analysis Code (NPAC) can predict the on- and off-design performance of axisymmetric or two-dimensional convergent and convergent-divergent nozzle geometries. NPAC accounts for the effects of overexpansion or underexpansion, flow divergence, wall friction, heat transfer, and small mass addition or loss across surfaces when the nozzle gross thrust and gross thrust coefficient are being computed. NPAC can be used to predict the performance of a given nozzle design or to develop a preliminary nozzle system design for subsequent analysis. The input required by NPAC consists of a simple geometry definition of the nozzle surfaces, the location of key nozzle stations (entrance, throat, exit), and the nozzle entrance flow properties. NPAC performs three analysis &qu...
A series of experiments were performed to investigate the effects of Mach number variation on the... more A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from...
Methods The main aspect of this project is to determine the feasibility of creating a hyper spect... more Methods The main aspect of this project is to determine the feasibility of creating a hyper spectral instrument on a miniaturized scale. A Cubesat unit is 10cm x 10cm x 10cm. This is the goal size of the instrument while still gaining the desired spectral and spatial resolution. Several Cubesat units can be ajoined to allow for larger payload, but these compartments would just be used for attitude control and power supply.
A Mathematical theory is developed to perform the calculations necessary to determine the wave dr... more A Mathematical theory is developed to perform the calculations necessary to determine the wave drag for slender bodies of non-circular cross section. The derivations presented in this report are based on extensions to supersonic linearized small perturbation theory. A numerical scheme is presented utilizing Fourier decomposition to compute the pressure coefficient on and about a slender body of arbitrary cross section.
24th Joint Propulsion Conference, 1988
For Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic thrQugh-fl... more For Mach 3.20 cruise propulsion systems, preliminary design studies for two supersonic thrQugh-flow fan primary inlets and a single core inlet were undertaken. Method of characteristics and one-dimensional performance techniques were applied to assess the potential improvements supersonic through-flow fan technology has over more conventional systems. A fixed geometry supersonic through-flow fan primary inlet was found to have better performance than a conventional inlet design on the basis of total pressure recovery, air flow, aerodynamic drags and size and weight. NOMENCLATURE A area BPR bypass ratio C D drag coefficient M Mach number
A series of experiments were performed to investigate the effects of Mach number variation on the... more A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from the dynamic pressure transducers were found to collapse to self-similar distributions as a function of the distance perpendicular to the separation line. The standard deviation of the pressure signals showed initial peaked distribution, with the maximum standard deviation point corresponding to the location of the separation line at Mach number 3.0 to 5.0. At Mach 2.0 the maximum standard deviation point was found to occur significantly upstream of the separation line. The intermittency distributions of the separation shock wave motion were found to be self-similar profiles for all Mach numbers. The intermittent region length was found to increase with Mach number and decrease with interaction sweepback angle. For Mach numbers 3.0 to 5.0 the separation line was found to correspond to high intermittencies or equivalently to the downstream locus of the separation shock wave motion. The Mach 2.0 tests, however, showed that the intermittent region occurs significantly upstream of the separation line. Power spectral densities measured in the intermittent regions were found to have self-similar frequency distributions when compared as functions of a Strouhal number for all Mach numbers and interaction sweepback angles. The maximum zero-crossing frequencies were found to correspond with the peak frequencies in the power spectra measured in the intermittent region.
A simple and accurate nozzle performance analysis methodology has been developed. The geometry mo... more A simple and accurate nozzle performance analysis methodology has been developed. The geometry modeling requirements are minimal and very flexible, thus allowing rapid design evaluations. The solution techniques accurately couple: continuity, momentum, energy, state, and other relations which permit fast and accurate calculations of nozzle gross thrust. The control volume and internal flow analyses are capable of accounting for the effects of: over/under expansion, flow divergence, wall friction, heat transfer, and mass addition/loss across surfaces. The results from the nozzle performance methodology are shown to be in excellent agreement with experimental data for a variety of nozzle designs over a range of operating conditions. preliminary nozzle system design, and subsequent performance analyses. Nozzle performance is typically described by determining three quantities: accepted engine airflow, WT, gross thrust coefficient, Cm, and aerodynamic drag coefficient, Co. It is also very important to be able to characterize nozzle performance over the entire vehicle flight and engine operation range, not just at the nozzle design point. The methodology presented
51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015
51st AIAA/SAE/ASEE Joint Propulsion Conference, 2015
Aircraft Design and Operations Meeting, 1989
A study is undertaken to investigate the engine airframe integration effects for supersonic throu... more A study is undertaken to investigate the engine airframe integration effects for supersonic through-flow fan engines installed on a Mach 3.20 supersonic cruise vehicle. Six different supersonic throughflow fan engine installations covering the effects of engine size, nacelle contour, nacelle placement, and approximate bypass plume effects are presented. The different supersonic through-flow fan installations are compared with a conventional turbine bypass engine configuration on the same basic airframe. The supersonic throughflow fan engine integrations are shown to be comparable to the turbine bypass engine configuration on the basis of installed nacelle wave drag. The supersonic throughflow fan engine airframe integrated vehicles have superior aerodynamic performance on the basis of maximum liftto-drag ratio than the turbine bypass engine installation over the entire operating Mach number range from i.i0 to 3.20. When approximate bypass plume modeling is included, the supersonic through-flow fan engine configuration shows even larger improvements over the turbine bypass engine configuration. NOMENCLATURE BPR bypass ratio CD drag coefficient, D/gS CL lift coefficient, L/qS D drag L lift M Mach number q dynamic pressure S reference area Subscrip__ f friction w wave INTRODUCTION In perusing the technologies required for efficient long range supersonic cruise aircraft, NASA has sponsored a number of studies to identify suitable propulsion system concepts. In the past, conventional and variable cycle engines were considered the most likely candidates. With renewed *Supervisor, Aerospace Analysis Section, Member AIAA Another desirable aspect of the engine cycle is that the bypass ratio decreases with increasing flight Math number, providing higher cruise thrust. Reference 3 shows that much of the improved specific fuel consumption for supersonic through-flow fan engines results from potential improvements in installation efficiencies for long range supersonic cruise applications. The possible improvements in inlet performance with supersonic through-flow fan engines has been shown in reference 4. This study addresses the engine airframe integration characteristics for supersonic through-flow fan propulsion systems, six different supersonic through-flow fan engine installations are examined in this study. The effects of engine size, nacelle contouring, nacelle placement, and approximate bypass plume modeling are investigated. Additionally, the supersonic through-flow fan (STFF) nacelle installations are compared with a conventional turbine bypass engine (TBE) nacelle installation on the basis of installed wave drags and maximum lift to drag ratios for complete engine airframe the
51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2013