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NASA NTRS2003

Structure of the Outer Cusp and Sources of the Cusp Precipitation during Intervals of a Horizontal IMF

The cusp represents a place where the magnetosheath plasma can directly penetrate into the magnetosphere. Since the main transport processes are connected with merging of the interplanetary and magnetospheric field lines: the interplanetary magnetic field (IMF) Orientation plays a decisive role in the formation of the high-altitude cusp. The importance of the sign of the IMF B(sub Z) component for this process was suggested about 40 years ago and later it was documented by many experimental investigations. However, situations when IMF Bz is the major IMF component are rather rare. The structure of the cusp during periods of a small IMF B(sub Z) is generally unknown, probably due to the fully 3-D nature of the interaction. The present case study reveals the importance of horizontal IMF components on the global magnetospheric configuration as well as on small-scale processes at the cusp-magnetosheath interface. We have used simultaneous measurements of several spacecraft (ISTP program) operating in different regions of interplanetary space and two closely spaced satellites (INTERBALL-1/MAGION-4) crossing the cusp-magnetosheath boundary to show the connection between the short- and large-scale phenomena. In the northern hemisphere, observations suggest a presence of two spots of cusp-like precipitation supplied by reconnection occurring simultaneously in both hemispheres. A source of this bifurcation is the positive IMF B(sub y) component further enhanced by the field draping in the magnetosheath. This magnetic field component shifts the entry point far away from the local noon but in opposite sense in either hemisphere. The cusp represents a place where the magnetosheath plasma can directly

Berchem, Jean↗

NASA NTRS2004

DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments

The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolution, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Moss, James N.↗

NASA NTRS2004

DSMC Simulations of Hypersonic Flows With Shock Interactions and Validation With Experiments

The capabilities of a relatively new direct simulation Monte Carlo (DSMC) code are examined for the problem of hypersonic laminar shock/shock and shock/boundary layer interactions, where boundary layer separation is an important feature of the flow. Flow about two model configurations is considered, where both configurations (a biconic and a hollow cylinder-flare) have recent published experimental measurements. The computations are made by using the DS2V code of Bird, a general two-dimensional/axisymmetric time accurate code that incorporates many of the advances in DSMC over the past decade. The current focus is on flows produced in ground-based facilities at Mach 12 and 16 test conditions with nitrogen as the test gas and the test models at zero incidence. Results presented highlight the sensitivity of the calculations to grid resolutions, sensitivity to physical modeling parameters, and comparison with experimental measurements. Information is provided concerning the flow structure and surface results for the extent of separation, heating, pressure, and skin friction.

Moss, James N.↗

NASA NTRS1998

Control of Rate-Bounded Hybrid Systems with Liveness Specifications

In the present paper we examine the control problem for a class of composite hybrid machines (CHMs) that consist of concurrent operation (employing synchronous composition) of elementary hybrid machines (EHMs), that allows both signal sharing and event synchronization. A controller can then be coupled with the plant by means of synchronous composition. We confine our attention to controllers that interact with the system only through event synchronization. We present an initial investigation of synthesis of liveness controllers for hybrid machines. To this end we define open hybrid machines as systems that can interact with the environment through event synchronization and can be therefor be "driven" to their marked configuration by user (controller). Liveness specifications must be associated with timing constraints. We may require that for a specified time limit, every run reach a marked configuration within that time limit. Alternatively, a more relaxed specification may be that, for some (unspecified) global time bound, every run of the system reach a marked configuration within that time bound. Finally, the least restrictive liveness requirement is that every run reach a marked configuration within a finite time limit (but we do not insist on the existence of a global time bound for all runs).

Heymann, Michael↗

NASA NTRS2008

Interactive Inverse Design Optimization of Fuselage Shape for Low-Boom Supersonic Concepts

This paper introduces a tool called BOSS (Boom Optimization using Smoothest Shape modifications). BOSS utilizes interactive inverse design optimization to develop a fuselage shape that yields a low-boom aircraft configuration. A fundamental reason for developing BOSS is the need to generate feasible low-boom conceptual designs that are appropriate for further refinement using computational fluid dynamics (CFD) based preliminary design methods. BOSS was not developed to provide a numerical solution to the inverse design problem. Instead, BOSS was intended to help designers find the right configuration among an infinite number of possible configurations that are equally good using any numerical figure of merit. BOSS uses the smoothest shape modification strategy for modifying the fuselage radius distribution at 100 or more longitudinal locations to find a smooth fuselage shape that reduces the discrepancies between the design and target equivalent area distributions over any specified range of effective distance. For any given supersonic concept (with wing, fuselage, nacelles, tails, and/or canards), a designer can examine the differences between the design and target equivalent areas, decide which part of the design equivalent area curve needs to be modified, choose a desirable rate for the reduction of the discrepancies over the specified range, and select a parameter for smoothness control of the fuselage shape. BOSS will then generate a fuselage shape based on the designer's inputs in a matter of seconds. Using BOSS, within a few hours, a designer can either generate a realistic fuselage shape that yields a supersonic configuration with a low-boom ground signature or quickly eliminate any configuration that cannot achieve low-boom characteristics with fuselage shaping alone. A conceptual design case study is documented to demonstrate how BOSS can be used to develop a low-boom supersonic concept from a low-drag supersonic concept. The paper also contains a study on how perturbations in the equivalent area distribution affect the ground signature shape and how new target area distributions for low-boom signatures can be constructed using superposition of equivalent area distributions derived from the Seebass-George-Darden (SGD) theory.

Li, Wu↗

NASA NTRS2010

Performance and Stability Analyses of Rocket Thrust Chambers with Oxygen/Methane Propellants

Liquid rocket engines using oxygen and methane propellants are being considered by the National Aeronautics and Space Administration (NASA) for future in-space vehicles. This propellant combination has not been previously used in flight-qualified engine systems developed by NASA, so limited test data and analysis results are available at this stage of early development. As part of activities for the Propulsion and Cryogenic Advanced Development (PCAD) project funded under the Exploration Technology Development Program, the NASA Marshall Space Flight Center (MSFC) has been evaluating capability to model combustion performance and stability for oxygen and methane propellants. This activity has been proceeding for about two years and this paper is a summary of results to date. Hot-fire test results of oxygen/methane propellant rocket engine combustion devices for the modeling investigations have come from several sources, including multi-element injector tests with gaseous methane from the 1980s, single element tests with gaseous methane funded through the Constellation University Institutes Program, and multi-element injector tests with both gaseous and liquid methane conducted at the NASA MSFC funded by PCAD. For the latter, test results of both impinging and coaxial element injectors using liquid oxygen and liquid methane propellants are included. Configurations were modeled with two one-dimensional liquid rocket combustion analysis codes, the Rocket Combustor Interactive Design and Analysis code and the Coaxial Injector Combustion Model. Special effort was focused on how these codes can be used to model combustion and performance with oxygen/methane propellants a priori, and what anchoring or calibrating features need to be applied, improved or developed in the future. Low frequency combustion instability (chug) occurred, with frequencies ranging from 150 to 250 Hz, with several multi-element injectors with liquid/liquid propellants, and was modeled using techniques from Wenzel and Szuch. High-frequency combustion instability also occurred at the first tangential (1T) mode, at about 4500 Hz, with several multi-element injectors with liquid/liquid propellants. Analyses of the transverse mode instability were conducted by evaluating injector resonances and empirical methods developed by Hewitt.

Hulka, James R.↗

NASA NTRS2009

Tropical Cyclogenesis in a Tropical Wave Critical Layer: Easterly Waves

The development of tropical depressions within tropical waves over the Atlantic and eastern Pacific is usually preceded by a "surface low along the wave" as if to suggest a hybrid wave-vortex structure in which flow streamlines not only undulate with the waves, but form a closed circulation in the lower troposphere surrounding the low. This structure, equatorward of the easterly jet axis, is identified herein as the familiar critical layer of waves in shear flow, a flow configuration which arguably provides the simplest conceptual framework for tropical cyclogenesis resulting from tropical waves, their interaction with the mean flow, and with diabatic processes associated with deep moist convection. The recirculating Kelvin cat's eye within the critical layer represents a sweet spot for tropical cyclogenesis in which a proto-vortex may form and grow within its parent wave. A common location for storm development is given by the intersection of the wave's critical latitude and trough axis at the center of the cat's eye, with analyzed vorticity centroid nearby. The wave and vortex live together for a time, and initially propagate at approximately the same speed. In most cases this coupled propagation continues for a few days after a tropical depression is identified. For easterly waves, as the name suggests, the propagation is westward. It is shown that in order to visualize optimally the associated Lagrangian motions, one should view the flow streamlines, or stream function, in a frame of reference translating horizontally with the phase propagation of the parent wave. In this co-moving frame, streamlines are approximately equivalent to particle trajectories. The closed circulation is quasi-stationary, and a dividing streamline separates air within the cat's eye from air outside.

Dunkerton, T. J.↗

NASA NTRS2014

Modeling and Prediction of the Noise from Non-Axisymmetric Jets

The new source model was combined with the original sound propagation model developed for rectangular jets to produce a new version of the rectangular jet noise prediction code. This code was validated using a set of rectangular nozzles whose geometries were specified by NASA. Nozzles of aspect ratios two, four and eight were studied at jet exit Mach numbers of 0.5, 0.7 and 0.9, for a total of nine cases. Reynolds-averaged Navier-Stokes solutions for these jets were provided to the contactor for use as input to the code. Quantitative comparisons of the predicted azimuthal and polar directivity of the acoustic spectrum were made with experimental data provided by NASA. The results of these comparisons, along with a documentation of the propagation and source models, were reported in a journal article publication (Ref. 4). The complete set of computer codes and computational modules that make up the prediction scheme, along with a user's guide describing their use and example test cases, was provided to NASA as a deliverable of this task. The use of conformal mapping, along with simplified modeling of the mean flow field, for noise propagation modeling was explored for other nozzle geometries, to support the task milestone of developing methods which are applicable to other geometries and flow conditions of interest to NASA. A model to represent twin round jets using this approach was formulated and implemented. A general approach to solving the equations governing sound propagation in a locally parallel nonaxisymmetric jet was developed and implemented, in aid of the tasks and milestones charged with selecting more exact numerical methods for modeling sound propagation, and developing methods that have application to other nozzle geometries. The method is based on expansion of both the mean-flowdependent coefficients in the governing equation and the Green's function in series of orthogonal functions. The method was coded and tested on two analytically prescribed mean flows which were meant to represent noise reduction concepts being considered by NASA. Testing (Ref. 5) showed that the method was feasible for the types of mean flows of interest in jet noise applications. Subsequently, this method was further developed to allow use of mean flow profiles obtained from a Reynolds-averaged Navier-Stokes (RANS) solution of the flow. Preliminary testing of the generalized code was among the last tasks completed under this contract. The stringent noise-reduction goals of NASA's Fundamental Aeronautics Program suggest that, in addition to potentially complex exhaust nozzle geometries, next generation aircraft will also involve tighter integration of the engine with the airframe. Therefore, noise generated and propagated by jet flows in the vicinity of solid surfaces is expected to be quite significant, and reduced-order noise prediction tools will be needed that can deal with such geometries. One important source of noise is that generated by the interaction of a turbulent jet with the edge of a solid surface (edge noise). Such noise is generated, for example, by the passing of the engine exhaust over a shielding surface, such as a wing. Work under this task supported an effort to develop a RANS-based prediction code for edge noise based on an extension of the classical Rapid Distortion Theory (RDT) to transversely sheared base flows (Refs. 6 and 7). The RDT-based theoretical analysis was applied to the generic problem of a turbulent jet interacting with the trailing edge of a flat plate. A code was written to evaluate the formula derived for the spectrum of the noise produced by this interaction and results were compared with data taken at NASA Glenn for a variety of jet/plate configurations and flow conditions (Ref. 8). A longer-term goal of this task was to work toward the development of a high-fidelity model of sound propagation in spatially developing non-axisymmetric jets using direct numerical methods for solving the relevant equations. Working with NASA Glenn Acoustics Branch personnel, numerical methods and boundary conditions appropriate for use in a high-resolution calculation of the full equations governing sound propagation in a steady base flow were identified. Computer codes were then written (by NASA) and tested (by OAI) for an increasingly complex set of flow conditions to validate the methods. The NASA-supplied codes were ported to the High-End Computing resources of the NASA Advanced Supercomputing facility for testing and validation against analytical (where possible) and independent numerical solutions. The cases which were completed during the course of this contract were solutions of the two-dimensional linearized Euler equations with no mean flow, a uniform mean flow and a nonuniform mean flow representative of a parallel flow jet.

Leib, Stewart J.↗

NASA NTRSDate not supplied

CFD Simulations of the Truncated IHF Semi-Elliptical Nozzle Flow: Convective Calibration Plate

This paper reports computational simulations of a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and flowfield over panel test articles. Both laminar and turbulent simulations are performed, and the computational results are compared with the calibration plate measurements. Comparisons of computed and measured heat fluxes indicate that the boundary layer in the semi-elliptical nozzle is transitional at certain conditions (relatively higher arc-heater pressures or mass flow rates).

Arc-jets↗

NASA NTRSDate not supplied

CFD Simulations of the Truncated IHF Semi-Elliptical Nozzle Flow: Convective Calibration Plate

This paper reports computational simulations of a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and the flowfields over panel test articles. Both laminar and turbulent simulations are performed, and the computational results are compared with the calibration plate measurements. Comparisons of computed and measured heat fluxes indicate that the boundary layer in the semi-elliptical nozzle is transitional at certain conditions (relatively higher arc-heater pressures or mass flow rates).

Arc-jets↗

NASA NTRSDate not supplied

On Characterization of Laminar to Turbulent Transitional Flow in the Truncated IHF Semi-Elliptical Nozzle

This paper reports computational analysis and experimental data to characterize the laminar to turbulent flow transition in a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and the flowfields over panel test articles, and comparisons with the measured test data. Both laminar and turbulent simulations are performed. Constant current studies are performed: the measured data are grouped at the same arc current levels with varying mass flow rates, and they are compared with corresponding computations. For all currents studied, these studies indicate that the semi-elliptical nozzle flow is transitional at relatively higher mass flow rates within the facility operating envelope.

Arc-jets↗

NASA NTRSDate not supplied

On Characterization of Laminar to Turbulent Transitional Flow in the Truncated IHF Semi-Elliptical Nozzle

This paper reports computational analysis and experimental data to characterize the laminar to turbulent flow transition in a high-enthalpy panel test facility at NASA Ames Research Center. The recently developed facility configuration uses a truncated version of an existing semi-elliptical nozzle in the 60-MW Interaction Heating Facility, and it provides testing capability for panel test articles as large as 43.2 cm x 43.2 cm (17 in x 17 in). A water-cooled calibration plate, which is attached to the semi-elliptical nozzle bottom surface and instrumented with twenty Gardon-gage heat flux calorimeters and twelve pressure gages, is used to characterize the convective heating environment over the panel test articles. The present analysis comprises computational fluid dynamics simulations of the nonequilibrium flowfield in the facility nozzle and test box, and the flowfields over panel test articles, and comparisons with the measured test data. Both laminar and turbulent simulations are performed. Constant current studies are conducted: the measured data are grouped at the same arc current levels with varying mass flow rates, and they are compared with corresponding computations. For all currents considered, these studies indicate that the semi-elliptical nozzle flow is transitional at relatively higher mass flow rates within the facility operating envelope. An approximate facility map for the transitional flow characterization is developed, showing laminar and transitional to turbulent flow regions.

Arc-jets↗

NASA NTRSDate not supplied

Advanced Diagnostics for Megahertz Imaging of Mixing, Fuel Spray, and Combustion Processes for Rotating Detonation Combustors

Recent advancements in megahertz rate, high-power, burst-mode laser technology are leveraged to perform and explore imaging measurements that spatially and temporally resolve the mixing, combustion, and detonation flow field in two laboratory-scale rotating detonation combustors (RDCs). In a non-premixed annular RDC, multiple imaging diagnostics are explored to investigate gaseous and liquid injector behavior, the detonation wave structure, and the propellant refill. In one instance, OH planar laser-induced fluorescence (OH-PLIF) imaging is performed up to a 2 MHz repetition rate to track the combustion products and reaction zone locations. In the same annular RDC, a single liquid fuel jet is injected, and laser-based 355-nm imaging of the fuel spray is performed up to a 1 MHz repetition rate. For this configuration, the annular RDC is used as a detonation driver to impose periodic detonation waves to interact with the fuel spray. Moreover, in this annular RDC, a range of tracer-based laser imaging measurements are explored to time-resolve the unsteady oxidizer air recovery and refill process. In a non-premixed linear RDC, planar imaging measurements of the fuel mixing are performed up to a 200 kHz repetition rate using PLIF of a tracer in the fuel supply. The fuel mixing imaging helps explain the origin of the observed pre and post wave burning, detonation structure, and enables quantifying injector recovery timescales. This paper will provide a high-level broad survey of diagnostics applied in these RDCs, lessons learned, and interesting observations.

Combustion↗

NASA NTRSDate not supplied

Advanced Diagnostics For Megahertz Imaging Of Mixing, Fuel Spray, And Combustion Processes For Rotating Detonation Combustors

Combustion↗

NASA NTRS1987

On the nonlinear aerodynamic and stability characteristics of a generic chine-forebody slender-wing fighter configuration

An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.

Erickson, Gary E.↗

NASA NTRS2019

Space Launch System Aeroacoustic Wind Tunnel Test Results

Characterization of accurate launch vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. For the National Aeronautics and Space Administration (NASA) Space Launch System (SLS), aeroacoustic environments have been derived primarily through sub-scale wind tunnel testing. Both optical techniques and high frequency pressure measurements have been utilized across multiple testing facilities and numerous vehicle configurations to develop a range of preliminary and detailed environments. As the vehicle has matured and evolved, the data collected from each subsequent configuration has allowed for comparison studies which isolate the effects of certain outer mold line (OML) features on measured fluctuating pressure levels. This paper presents observations on some of those effects for features which include abort system protuberances, various fairings geometries, interstage flanges, and multibody interactions between a central core and fall away boosters. These features, and the flow conditions produced by them, are broadly applicable to many launch vehicle configurations.

Steva, Thomas B.↗

NASA NTRS2014

Flow-Field Investigation of Gear-Flap Interaction on a Gulfstream Aircraft Model

Off-surface flow measurements of a high-fidelity 18% scale Gulfstream aircraft model in landing configuration with the main landing gear deployed are presented. Particle Image Velocimetry (PIV) and Laser Velocimetry (LV) were used to measure instantaneous velocities in the immediate vicinity of the main landing gear and its wake and near the inboard tip of the flap. These measurements were made during the third entry of a series of tests conducted in the NASA Langley Research Center (LaRC) 14- by 22-Foot Subsonic Tunnel (14 x 22) to obtain a comprehensive set of aeroacoustic measurements consisting of both aerodynamic and acoustic data. The majority of the off-body measurements were obtained at a freestream Mach number of 0.2, angle of attack of 3 degrees, and flap deflection angle of 39 degrees with the landing gear on. A limited amount of data was acquired with the landing gear off. LV was used to measure the velocity field in two planes upstream of the landing gear and to measure two velocity profiles in the landing gear wake. Stereo and 2-D PIV were used to measure the velocity field over a region extending from upstream of the landing gear to downstream of the flap trailing edge. Using a special traverse system installed under the tunnel floor, the velocity field was measured at 92 locations to obtain a comprehensive picture of the pertinent flow features and characteristics. The results clearly show distinct structures in the wake that can be associated with specific components on the landing gear and give insight into how the wake is entrained by the vortex at the inboard tip of the flap.

Yao, Chung-Sheng↗

NASA NTRSDate not supplied

Comparing RotCFD Predictions of the Multirotor Test Bed with Experimental Results

The Multirotor Test Bed (MTB) is a new capability for testing a wide array of advanced vertical take-off and landing (VTOL) rotor configurations, with a primary focus on testing in the U.S. Army 7- by 10-ft Wind Tunnel at NASA Ames Research Center. The MTB was designed to allow adjustment of the vertical, lateral, and longitudinal placement of each rotor, as well as allow tilt adjustment of each rotor and pitch adjustment of the whole assembly. Each rotor can tilt forward 90 deg and backwards 5 deg. In addition, the entire MTB can tilt forward 20 deg and backwards 10 deg. This flexibility allows the system to be tested in many different configurations. There is a six-axis load cell under each rotor assembly, to measure both the steady and dynamic loads produced by each rotor. The wind tunnel scales can measure loads on the full assembly. The overall goal of the MTB project is to help gain a better understanding of the performance, control, interactional aerodynamics, and acoustics of multirotor systems. A hybrid CFD tool called RotCFD (Rotorcraft Computational Fluid Dynamics) was used to simulate the MTB in several testing configurations. This paper explains the method of running the RotCFD simulations and explores the results from the simulations. The objective of this paper is to compare the RotCFD simulation results with the MTB wind tunnel test data, seeking to further validate RotCFD for multirotor systems and assess the influence of aerodynamic interactions on individual rotor performance.

RotCFD↗