The Experts below are selected from a list of 63195 Experts worldwide ranked by ideXlab platform

Spivey, Natalie D. - One of the best experts on this subject based on the ideXlab platform.

  • High-Temperature Modal Survey of a Hot-Structure Control Surface
    2011
    Co-Authors: Spivey, Natalie D.
    Abstract:

    Ground vibration tests are routinely conducted for supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicles, thermoelastic vibration testing techniques are neither well established nor routinely performed. New high-temperature material systems, fabrication technologies and high-temperature sensors expand the opportunities to develop advanced techniques for performing ground vibration tests at elevated temperatures. When high-temperature materials, which increase in stiffness when heated, are incorporated into a hot-structure that contains metallic components that decrease in stiffness when heated, the interaction between those materials can affect the hypersonic flutter analysis. A high-temperature modal survey will expand the research database for Hypersonics and improve the understanding of this dual-material interaction. This report discusses the vibration testing of the carbon-silicon carbide Ruddervator Subcomponent Test Article, which is a truncated version of a full-scale hot-structure control surface. Two series of room-temperature modal test configurations were performed in order to define the modal characteristics of the test article during the elevated-temperature modal survey: one with the test article suspended from a bungee cord (free-free) and the second with it mounted on the strongback (fixed boundary). Testing was performed in the NASA Dryden Flight Research Center Flight Loads Laboratory Large Nitrogen Test Chamber

  • High-Temperature Modal Survey of a Hot-Structure Control Surface
    2009
    Co-Authors: Spivey, Natalie D.
    Abstract:

    NASA Dryden is directing a program to test a Carbon-Silicon Carbide (C/SiC) Ruddervator Subcomponent Test Article (RSTA). The RSTA is a truncated version of the full-scale X-37 hot-structure control surface incorporating all major features including the metallic spindle and five major C/SiC quasi isotropic lay-up components secured with C/SiC fasteners. As part of NASA's Aeronautics Research Mission Directorate Hypersonics program, the RSTA will undergo thermal-structural testing to develop an extensive database for future structural design and analysis methodology validation. Ground Vibration Tests (GVTs) are routinely conducted for model validation in supporting flutter analysis for subsonic and supersonic vehicles; however, for hypersonic vehicle applications, GVT techniques are not well-established. New fabrication technologies, high-temperature materials systems, and sensors offer new opportunities to develop techniques for performing GVTs at elevated temperatures. The RSTA is comprised of materials whose stiffness both increases and decreases with increasing temperature. The impact of this type of material system interaction must be understood as it will ultimately affect hypersonic flutter analysis. The test objectives are to perform room-temperature GVTs, develop the capability to conduct high-temperature GVTs and to compare and generate an understanding of the modal characteristics which capture RSTA's material interaction when subjected to temperature varying conditions

Leidy, Andrew Norbert - One of the best experts on this subject based on the ideXlab platform.

  • An Experimental Characterization of 3-D Transitional Shock Wave Boundary Layer Interactions at Mach 6
    2019
    Co-Authors: Leidy, Andrew Norbert
    Abstract:

    Hypersonics is of current national interest, but improved understanding of fundamental flow physics is required for safe and efficient vehicle design. The objective of the present study was to expand the knowledgebase of shock wave boundary layer interaction flows during transition at a high Mach number since these interactions are likely to occur on hypersonic aircraft. The approach was to experimentally determine how the dynamics of the shock structure, the fluctuations within it, and the resulting thermal and acoustic loads, change as the flow evolves through its transitional regime. Tests were conducted on a canonical cylinder-induced 3-D shock wave boundary layer interaction geometry at Mach 5.8 in the Actively Controlled Expansion hypersonic wind tunnel. The model was tested in different configurations to isolate the effects of the boundary layer trips and the shock generator. The interaction excited a 40 kHz (possibly second mode) in-stability, causing transition just downstream of the separation shock. A transitional boundary layer was only achieved on the baseline model with trips at Re=7M/m, which demonstrated that a transitional incoming boundary layer is not required to produce a transition interaction. Time-resolved schlieren imaging revealed disturbances emerging from the supersonic jet and ascending the cylinder with a characteristic frequency near 20 kHz. The separation shock motion frequency was O(1 kHz) and was fed by disturbances originating near the base of the cylinder. The film coefficient was found to be the heat transfer parameter of interest since it scaled roughly linearly with Reynolds number. It revealed fundamental differences in heating at the reattachment arc for configurations with and without trips and indicated higher heating in that region for a laminar SBLI. Cylinder sweep had an impact on fluctuation levels and thermal loads. Sweeping the cylinder back 15 degrees significantly reduced the extent of the interaction and dropped the RMS pressure fluctuations and heating loads at the base of the cylinder by roughly 50%. Alternatively, sweeping the cylinder forward 15 degrees led to fluctuations on the order of the freestream static pressure and the highest heating levels observed for this campaign

Masatomi Nishio - One of the best experts on this subject based on the ideXlab platform.

  • Methods for visualizing hypersonic shock-wave/boundary-layer interaction using electrical discharges
    AIAA Journal, 1996
    Co-Authors: Masatomi Nishio
    Abstract:

    Two methods for visualizing the spatial flowfield of hypersonic shock-wave/boundary-layer interaction were developed by utilizing the radiation of electrical discharges. One method visualizes boundary layers in hypersonic flow, and the other visualizes streamlines near wall surfaces in hypersonic flow. These two methods were applied to the visualization of hypersonic shock-wave/boundary-layer interaction, and the results confirmed that the two methods are useful for visualizing the spatial flowfield of hypersonic shock-wave/boundary-layer interaction. The experiments were carried out using a hypersonic shock tunnel, with Mach number of 10 and duration of 10 m.s.

Saunders J. D. - One of the best experts on this subject based on the ideXlab platform.

  • Testing of the NASA Hypersonics Project Combined Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LlMX)
    2012
    Co-Authors: Weir L. J., Thomas, Scott R., Suder K. L., Sanders B. W., Stueber T. J., Saunders J. D.
    Abstract:

    Status on an effort to develop Turbine Based Combined Cycle (TBCC) propulsion is described. This propulsion technology can enable reliable and reusable space launch systems. TBCC propulsion offers improved performance and safety over rocket propulsion. The potential to realize aircraft-like operations and reduced maintenance are additional benefits. Among most the critical TBCC enabling technologies are: 1) mode transition from turbine to scramjet propulsion, 2) high Mach turbine engines and 3) TBCC integration. To address these TBCC challenges, the effort is centered on a propulsion mode transition experiment and includes analytical research. The test program, the Combined-Cycle Engine Large Scale Inlet Mode Transition Experiment (CCE LIMX), was conceived to integrate TBCC propulsion with proposed hypersonic vehicles. The goals address: (1) dual inlet operability and performance, (2) mode-transition sequences enabling a switch between turbine and scramjet flow paths, and (3) turbine engine transients during transition. Four test phases are planned from which a database can be used to both validate design and analysis codes and characterize operability and integration issues for TBCC propulsion. In this paper we discuss the research objectives, features of the CCE hardware and test plans, and status of the parametric inlet characterization testing which began in 2011. This effort is sponsored by the NASA Fundamental Aeronautics Hypersonics projec

Juntao Chang - One of the best experts on this subject based on the ideXlab platform.

  • Hysteresis phenomenon of hypersonic inlet at high Mach number
    Acta Astronautica, 2016
    Co-Authors: Xiaoliang Jiao, Juntao Chang, Zhongqi Wang
    Abstract:

    Abstract When the hypersonic inlet works at a Mach number higher than the design value, the hypersonic inlet is started with a regular reflection of the external compression shock at the cowl, whereas a Mach reflection will result in the shock propagating forwards to cause a shock detachment at the cowl lip, which is called “local unstart of inlet”. As there are two operation modes of hypersonic inlet at high Mach number, the mode transition may occur with the operation condition of hypersonic inlet changing. A cowl-angle-variation-induced hysteresis and a downstream-pressure-variation-induced hysteresis in the hypersonic inlet start↔local unstart transition are obtained by viscous numerical simulations in this paper. The interaction of the external compression shock and boundary layer on the cowl plays a key role in the hysteresis phenomenon. Affected by the transition of external compression shock reflection at the cowl and the transition between separated and attached flow on the cowl, a hysteresis exists in the hypersonic inlet start↔local unstart transition. The hysteresis makes the operation of a hypersonic inlet very difficult to control. In order to avoid hysteresis phenomenon and keep the hypersonic inlet operating in a started mode, the control route should never pass through the local unstarted boundary.

  • Effects of boundary-layer bleeding on unstart oscillatory flow of hypersonic inlets
    Aeronautical Journal, 2010
    Co-Authors: Juntao Chang, D. Yu
    Abstract:

    The unsteady flowfield of a series of mixed-compression hypersonic inlets with different bleeding rates were numerically simulated. Firstly unstart oscillatory flow of hypersonic inlets caused by downstream massflow choking was discussed. Then the effects of boundary layer bleeding on the averaged performance parameter of hypersonic inlets, and on the dominant amplitude and frequency of unstart oscillatory flow of hypersonic inlets were presented. The reasons why the boundary-layer bleeding can suppress unstart oscillatory flow of hypersonic inlets were analysed. In conclusion, the averaged performance parameter of hypersonic inlets during a big buzz is improved greatly, and the dominant frequency of unstart oscillatory flow of hypersonic inlets is reduced in contrast with no bleeding, and all these are benefit to the design and operation of hypersonic inlets.

  • Effects of wall cooling on performance parameters of hypersonic inlets
    Acta Astronautica, 2009
    Co-Authors: Juntao Chang, Wen Bao, Yi Fan, Yi Shen
    Abstract:

    Abstract The internal flowfield of a 2-D mixed-compression hypersonic inlet was simulated numerically at different wall temperatures. The effects of wall cooling on performance parameters of the hypersonic inlet without and with the backpressure, and especially on the maximum backpressure of a fixed-geometry hypersonic inlet were discussed. The inner physical mechanism of wall cooling improving the maximum backpressure of hypersonic inlets was analyzed. In contrast with no wall cooling, the static pressure ratio of hypersonic inlets without backpressure is reduced slightly, but the mass-captured coefficient, total-pressure recovery coefficient and the flow uniformity of hypersonic inlets at the isolator exit are improved by the action of wall cooling. The interaction between boundary layers and shocks is weakened due to wall cooling, which leads to that the boundary layers separations at the entrance of the isolator caused by the high backpressure occur later, and it can improve the maximum backpressure ratio of hypersonic inlets. With the wall temperatures decreasing, the maximum backpressure ratio and mass-captured coefficient are added, and the total-pressure recovery coefficient of hypersonic inlets is reduced.

  • Hypersonic inlet control with pulse periodic energy addition
    Proceedings of the Institution of Mechanical Engineers Part G: Journal of Aerospace Engineering, 2009
    Co-Authors: Juntao Chang, Wen Bao, Yi Fan, Yi Shen, Weixing Zhou
    Abstract:

    AbstractThe theory of pulse periodic energy addition to hypersonic airflow off a vehicle to increase air mass capture and improve the performance of hypersonic inlets at Mach numbers below the design value is explored. The unsteady numerical simulation of hypersonic inlets with energy addition was performed, and pulse periodic energy addition was introduced to discuss the possibility of the steady performance improvement of hypersonic inlets. The numerical simulation of hypersonic inlets at different pulse periodic energy addition rates was carried out. The simulation result shows that the performance parameters are improved at pulse periodic than at fixed values for the same averaged energy addition, and the energy addition rate needed is less at pulse periodic than at fixed values for the same performance parameters of hypersonic inlets.