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  • Other Sources  (130)
  • Fluid Mechanics and Thermodynamics  (119)
  • Aircraft Stability and Control
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  • 2015-2019  (130)
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  • 2016  (130)
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  • 2015-2019  (130)
  • 1995-1999
  • 1955-1959
  • 1950-1954
  • 1945-1949
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  • 1
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    Orion Exploration Flight Test Reaction Control System Jet Interaction Heating Environment from Flight Data (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-19
    Description: The Orion Multi-Purpose Crew Vehicle (MPCV) Reaction Control System (RCS) is critical to guide the vehicle along the desired trajectory during re-entry. However, this system has a significant impact on the convective heating environment to the spacecraft. Heating augmentation from the jet interaction (JI) drives thermal protection system (TPS) material selection and thickness requirements for the spacecraft. This paper describes the heating environment from the RCS on the afterbody of the Orion MPCV during Orion's first flight test, Exploration Flight Test 1 (EFT-1). These jet plumes interact with the wake of the crew capsule and cause an increase in the convective heating environment. Not only is there widespread influence from the jet banks, there may also be very localized effects. The firing history during EFT-1 will be summarized to assess which jet bank interaction was measured during flight. Heating augmentation factors derived from the reconstructed flight data will be presented. Furthermore, flight instrumentation across the afterbody provides the highest spatial resolution of the region of influence of the individual jet banks of any spacecraft yet flown. This distribution of heating augmentation across the afterbody will be derived from the flight data. Additionally, trends with possible correlating parameters will be investigated to assist future designs and ground testing programs. Finally, the challenges of measuring JI, applying this data to future flights and lessons learned will be discussed.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: JSC-CN-34678 , AIAA Thermophysics Conference; Jun 13, 2016 - Jun 17, 2016; Washington, DC; United States
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  • 2
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    Boundary Layer Transition During the Orion Exploration Flight Test 1 (EFT-1) (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-19
    Description: Boundary layer transition was observed in the thermocouple data on the windside backshell of the Orion reentry capsule. Sensors along the windside centerline, as well as off-centerline, indicated transition late in the flight at approximately Mach 4 conditions. Transition progressed as expected, beginning at the sensors closest to the forward bay cover (FBC) and moving towards the heatshield. Sensors placed in off-centerline locations did not follow streamlines, so the progression of transition observed in these sensors is less intuitive. Future analysis will include comparisons to pre-flight predictions and expected transitional behavior will be investigated. Sensors located within the centerline and off-centerline launch abort system (LAS) attach well cavities on the FBC also showed indications of boundary layer transition. The transition within the centerline cavity was observed in the temperature traces prior to transition onset on the sensors upstream of the cavity. Transition behavior within the off centerline LAS attach well cavity will also be investigated. Heatshield thermocouples were placed within Avcoat plugs to attempt to capture transitional behavior as well as better understand the aerothermal environments. Thermocouples were placed in stacks of two or five vertically within the plugs, but the temperature data obtained at the sensors closest to the surface did not immediately indicate transitional behavior. Efforts to use the in depth thermocouple temperatures to reconstruct the surface heat flux are ongoing and any results showing the onset of boundary layer transition obtained from those reconstructions will also be included in this paper. Transition on additional features of interest, including compression pad ramps, will be included if it becomes available.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: JSC-CN-34666 , AIAA Thermophysics Conference; Jun 13, 2016 - Jun 17, 2016; Washington, DC; United States
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  • 3
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    Experimental Investigation of Subsonic Turbulent Boundary Layer Flow Over a Wall-Mounted Axisymmetric Hill (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An important goal for modern fluid mechanics experiments is to provide datasets which present a challenge for Computational Fluid Dynamics simulations to reproduce. Such "CFD validation experiments" should be well-characterized and well-documented, and should investigate flows which are difficult for CFD to calculate. It is also often convenient for the experiment to be challenging for CFD in some aspects while simple in others. This report is part of the continuing documentation of a series of experiments conducted to characterize the flow around an axisymmetric, modified-cosine-shaped, wall-mounted hill named "FAITH" (Fundamental Aero Investigates The Hill). Computation of this flow is easy in some ways - subsonic flow over a simple shape - while being complex in others - separated flow and boundary layer interactions. The primary set of experiments were performed on a 15.2 cm high, 45.7 cm base diameter machined aluminum model that was tested at mean speeds of 50 m/s (Reynolds Number based on height = 500,000). The ratio of model height to boundary later height was approximately 3. The flow was characterized using surface oil flow visualization, Cobra probe to determine point-wise steady and unsteady 3D velocities, Particle Image Velocimetry (PIV) to determine 3D velocities and turbulence statistics along specified planes, Pressure Sensitive Paint (PSP) to determine mean surface pressures, and Fringe Imaging Skin Friction (FISF) to determine surface skin friction magnitude and direction. A set of pathfinder experiments were also performed in a water channel on a smaller scale (5.1 cm high, 15.2 cm base diameter) sintered nylon model. The water channel test was conducted at a mean test section speed of 3 cm/s (Reynolds Number of 1500), but at the same ratio of model height to boundary layer thickness. Dye injection from both the model and an upstream rake was used to visualize the flow. This report summarizes the experimental set-up, techniques used, and data acquired. It also describes some details of the dataset that is being constructed for use by other researchers, especially the CFD community.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: STO-MP-AVT-246 , ARC-E-DAA-TN35029 , ARC-E-DAA-TN34022 , NATO CSO AVT-246 Specialists Meeting; Sep 26, 2016 - Sep 28, 2016; Avila; Spain
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  • 4
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    Pressure Loss Predictions of the Reactor Simulator Subsystem at NASA Glenn Research Center (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Testing of the Fission Power System (FPS) Technology Demonstration Unit (TDU) is being conducted at NASA Glenn Research Center. The TDU consists of three subsystems: the reactor simulator (RxSim), the Stirling Power Conversion Unit (PCU), and the heat exchanger manifold (HXM). An annular linear induction pump (ALIP) is used to drive the working fluid. A preliminary version of the TDU system (which excludes the PCU for now) is referred to as the "RxSim subsystem" and was used to conduct flow tests in Vacuum Facility 6 (VF 6). In parallel, a computational model of the RxSim subsystem was created based on the computer-aided-design (CAD) model and was used to predict loop pressure losses over a range of mass flows. This was done to assess the ability of the pump to meet the design intent mass flow demand. Measured data indicates that the pump can produce 2.333 kg/sec of flow, which is enough to supply the RxSim subsystem with a nominal flow of 1.75 kg/sec. Computational predictions indicated that the pump could provide 2.157 kg/sec (using the Spalart-Allmaras (SA) turbulence model) and 2.223 kg/sec (using the k- turbulence model). The computational error of the predictions for the available mass flow is 0.176 kg/sec (with the S-A turbulence model) and -0.110 kg/sec (with the k- turbulence model) when compared to measured data.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA/TM-2016-218913 , AIAA Paper 2015-3906 , E-19169 , GRC-E-DAA-TN25966 , International Energy Conversion Engineering; Jul 27, 2015 - Jul 29, 2015; Orlando, FL; United States
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  • 5
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    Extension of the Viscous Collision Limiting Direct Simulation Monte Carlo Technique to Multiple Species (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: There are many flows fields that span a wide range of length scales where regions of both rarefied and continuum flow exist and neither direct simulation Monte Carlo (DSMC) nor computational fluid dynamics (CFD) provide the appropriate solution everywhere. Recently, a new viscous collision limited (VCL) DSMC technique was proposed to incorporate effects of physical diffusion into collision limiter calculations to make the low Knudsen number regime normally limited to CFD more tractable for an all-particle technique. This original work had been derived for a single species gas. The current work extends the VCL-DSMC technique to gases with multiple species. Similar derivations were performed to equate numerical and physical transport coefficients. However, a more rigorous treatment of determining the mixture viscosity is applied. In the original work, consideration was given to internal energy non-equilibrium, and this is also extended in the current work to chemical non-equilibrium.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-23668 , International Symposium on Rarified Gas Dynamics; Jul 10, 2016 - Jul 14, 2016; Victoria; Canada
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  • 6
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    CFD Analysis of Mixing Characteristics of Several Fuel Injectors at Hypervelocity Flow Conditions (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: CFD analysis is presented of the mixing characteristics and performance of three fuel injectors at hypervelocity flow conditions. The calculations were carried out using the VULCAN-CFD solver and Reynolds-Averaged Simulations (RAS). The high Mach number flow conditions match those proposed for the planned experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships relevant to flight Mach numbers greater than eight. Because of the high Mach number flow considered, the injectors consist of a fuel placement device, a strut; and a fluidic vortical mixer, a ramp. These devices accomplish the necessary task of distributing and mixing fuel into the supersonic cross-flow albeit via different strategies. Both of these devices were previously studied at lower flight Mach numbers where they exhibited promising performance in terms of mixing efficiency and total pressure recovery. For comparison, a flush-wall injector is also included. This type of injector generally represents the simplest method of introducing fuel into a scramjet combustor, however, at high flight Mach number conditions, the dynamic pressure needed to induce sufficient fuel penetration may be difficult to achieve along with other requirements such as achieving desired levels of fuel-to-air mixing at the required equivalence ratio. The three injectors represent the baseline configurations planned for the experiments. The current work discusses the mixing flow field behavior and differences among the three fuel injectors, mixing performance as described by the mixing efficiency and the total pressure recovery, and performance considerations based on the thrust potential.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-23413 , AIAA Propulsion and Energy Forum and Exposition; Jul 25, 2016 - Jul 27, 2016; Salt Lake, UT; United States
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  • 7
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    Modeling the Rapid Boil-Off of a Cryogenic Liquid When Injected into a Low Pressure Cavity (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Many launch vehicle cryogenic applications require the modeling of injecting a cryogenic liquid into a low pressure cavity. The difficulty of such analyses lies in accurately predicting the heat transfer coefficient between the cold liquid and a warm wall in a low pressure environment. The heat transfer coefficient and the behavior of the liquid is highly dependent on the mass flow rate into the cavity, the cavity wall temperature and the cavity volume. Testing was performed to correlate the modeling performed using Thermal Desktop and Sinda Fluint Thermal and Fluids Analysis Software. This presentation shall describe a methodology to model the cryogenic process using Sinda Fluint, a description of the cryogenic test set up, a description of the test procedure and how the model was correlated to match the test results.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: KSC-E-DAA-TN34363 , NASA Thermal Fluids & Analysis Workshop (TFAWS 2016); Aug 01, 2016 - Aug 05, 2016; Mountain View, CA; United States
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  • 8
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    Tapping the Brake for Entry, Descent, and Landing (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A matrix of simulations of hypersonic flow over blunt entry vehicles with steady and pulsing retropropulsion jets is presented. Retropropulsion in the supersonic domain is primarily designed to reduce vehicle velocity directly with thrust. Retropropulsion in the hypersonic domain may enable significant pressure recovery through unsteady, oblique shocks while providing a buffer of reactant gases with relatively low total temperature. Improved pressure recovery, a function of Mach number squared and oblique shock angle, could potentially serve to increase aerodynamic drag in this domain. Pulsing jets are studied to include an additional degree of freedom to search for resonances in an already unsteady flow domain with an objective to maximize the time-averaged drag coefficient. In this paradigm, small jets with minimal footprints of the nozzle exit on the vehicle forebody may be capable of delivering the requisite perturbations to the flow. Simulations are executed assuming inviscid, symmetric flow of a perfect gas to enable a rapid assessment of the parameter space (nozzle geometry, plenum conditions, jet pulse frequency). The pulsed-jet configuration produces moderately larger drag than the constant jet configuration but smaller drag than the jet-off case in this preliminary examination of a single design point. The fundamentals of a new algorithm for this challenging application with time dependent, interacting discontinuities using the feature detection capabilities of Walsh functions are introduced.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-22796 , AIAA Fluid Dynamics Conference; Jun 13, 2016 - Jun 17, 2016; Washington, DC; United States
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  • 9
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    Flight Experiment Verification of Shuttle Boundary Layer Transition Prediction Tool (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Boundary layer transition at hypersonic conditions is critical to the design of future high-speed aircraft and spacecraft. Accurate methods to predict transition would directly impact the aerothermodynamic environments used to size a hypersonic vehicle's thermal protection system. A transition prediction tool, based on wind tunnel derived discrete roughness correlations, was developed and implemented for the Space Shuttle return-to-flight program. This tool was also used to design a boundary layer transition flight experiment in order to assess correlation uncertainties, particularly with regard to high Mach-number transition and tunnel-to-flight scaling. A review is provided of the results obtained from the flight experiment in order to evaluate the transition prediction tool implemented for the Shuttle program.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-22834 , AIAA Aviation 2016; Jun 13, 2016 - Jun 17, 2016; Washington, DC; United States
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  • 10
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    Sweeping Jet Optimization Studies (2016)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Progress on experimental efforts to optimize sweeping jet actuators for active flow control (AFC) applications with large adverse pressure gradients is reported. Three sweeping jet actuator configurations, with the same orifice size but dierent internal geometries, were installed on the flap shoulder of an unswept, NACA 0015 semi-span wing to investigate how the output produced by a sweeping jet interacts with the separated flow and the mechanisms by which the flow separation is controlled. For this experiment, the flow separation was generated by deflecting the wing's 30% chord trailing edge flap to produce an adverse pressure gradient. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the three actuator configurations. The actuator with the largest jet deflection angle, at the pressure ratios investigated, was the most efficient at controlling flow separation on the flap of the model. Oil flow visualization studies revealed that the flow field controlled by the sweeping jets was more three-dimensional than expected. The results presented also show that the actuator spacing was appropriate for the pressure ratios examined.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NF1676L-22842 , 2016 AIAA Aviation Technology, Integration, and Operations Conference; Jun 13, 2016 - Jun 17, 2016; Washington, DC; United States
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