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  • 1
    Electronic Resource
    Electronic Resource
    Introduction to research on Hirschman's Exit, Voice, and Loyalty model (1992)
    Springer
    Employee responsibilities and rights journal 5 (1992), S. 187-190 
    Details
    ISSN: 1573-3378
    Keywords: exit, voice, and loyalty ; dissatisfaction model ; Hirschman's model ; controversy about Hirschman's model
    Source: Springer Online Journal Archives 1860-2000
    Topics: Economics
    Notes: Abstract This article introduces this special issue of theEmployee Responsibilities and Rights Journal on recent work exploring Albert O. Hirschman's Exit, Voice, and Loyalty model of dissatisfaction. This special issue provides a forum for researchers and theorists with various perspectives on the model to present their ideas in one place. There are six original articles in this issue, and one discussion piece. While these articles do not resolve the controversies surrounding the Exit, Voice, and Loyalty model, they do provide a clear picture of the current status of research and theory on dissatisfaction in organizations from this perspective.
    Type of Medium: Electronic Resource
    URL: http://dx.doi.org/10.1007/BF01385046
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    Paper (German National Licenses)
    Fulltext
  • 2
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    Current Status on Radiation Modeling for the Hayabusa Re-entry (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: On June 13, 2010 the Japanese Hayabusa capsule performed its reentry into the Earths atmosphere over Australia after a seven year journey to the asteroid Itokawa. The reentry was studied by numerous imaging and spectroscopic instruments onboard NASA's DC-8 Airborne Laboratory and from three sites on the ground, in order to measure surface and plasma radiation generated by the Hayabusa Sample Return Capsule (SRC). Post flight, the flow solutions were recomputed to include the whole flow field around the capsule at 11 points along the reentry trajectory using updated trajectory information. Again, material response was taken into account to obtain most reliable surface temperature information. These data will be used to compute thermal radiation of the glowing heat shield and plasma radiation by the shock/post-shock layer system to support analysis of the experimental observation data. For this purpose, lines of sight data are being extracted from the flow field volume grids and plasma radiation will be computed using NEQAIR [4] which is a line-by-line spectroscopic code with one-dimensional transport of radiation intensity. The procedures being used were already successfully applied to the analysis of the observation of the Stardust reentry [5].
    Keywords: Space Radiation
    Type: ARC-E-DAA-TN3212 , Workshop on Re-Entry Emission Signatures V; Mar 21, 2011 - Mar 22, 2011; Brisbane; Australia
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    InSight Aerothermal Environment Assessment (2020)
    In:  CASI
    Details
    Publication Date: 2019-08-01
    Description: The InSight spacecraft was proposed to be a build-to-print copy of the Phoenix vehicle due to the knowledge that the lander payload would be similar and the trajectory would be similar. However, the InSight aerothermal analysts, based on tests performed in CO2 during the Mars Science Laboratory mission (MSL) and completion of Russian databases, considered radiative heat flux to the aftbody from the wake for the first time for a US Mars mission. The combined convective and radiative heat flux was used to determine if the as-flown Phoenix thermal protection system (TPS) design would be sufficient for InSight. All analyses showed that the design would be adequate. Once the InSight lander was successfully delivered to Mars on November 26, 2018, work began to reconstruct the atmosphere and trajectory in order to evaluate the aerothermal environments that were actually encountered by the spacecraft and to compare them to the design environments.The best estimated trajectory (BET) reconstructed for the InSight atmospheric entry fell between the two trajectories considered for the design, when looking at the velocity versus altitude values. The maximum heat rate design trajectory (MHR) flew at a higher velocity and the maximum heat load design trajectory (MHL) flew at a lower velocity than the BET. For TPS sizing, the MHL trajectory drove the design. Reconstruction has shown that the BET flew for a shorter time than either of the design environments, hence total heat load on the vehicle should have been less than used in design. Utilizing the BET, both DPLR and LAURA were first run to analyze the convective heating on the vehicle with no angle of attack. Both codes were run with axisymmetric, laminar flow in radiative equilibrium and vibrational non-equilibrium with a surface emissivity of 0.8. Eight species Mitcheltree chemistry was assumed with CO2, CO, N2, O2, NO, C, N, and O. Both codes agreed within 1% on the forebody and had the expected differences on the aftbody. The NEQAIR and HARA codes were used to analyze the radiative heating on the vehicle using full spherical ray-tracing. The codes agreed within 5% on most aftbody points of interest.The LAURA code was then used to evaluate the conditions at angle of attack at the peak heating and peak pressure times. Boundary layer properties were investigated to confirm that the flow over the forebody was laminar for the flight.Comparisons of the aerothermal heating determined for the reconstructed trajectory to the design trajectories showed that the as-flown conditions were less severe than design
    Keywords: Fluid Mechanics and Thermodynamics
    Type: ARC-E-DAA-TN69598 , AIAA SciTech 2020; Jan 06, 2020 - Jan 10, 2020; Orlando, FL; United States
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    NASA TECHNICAL REPORTS
  • 4
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    Allowable Trajectory Variations for Space Shuttle Orbiter Entry-Aeroheating CFD (2008)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Reynolds-number criteria are developed for acceptable variations in Space Shuttle Orbiter entry trajectories for use in computational aeroheating analyses. The criteria determine if an existing computational fluid dynamics solution for a particular trajectory can be extrapolated to a different trajectory. The criteria development considers twelve types of computational aeroheating data, such as boundary layer thickness. For each type of datum, the allowable uncertainty contribution due to trajectory variation has been set by the Entry Aeroheating Subsystem team. Then Reynolds-number relations between trajectory variation and output uncertainty are determined. From these relations the criteria are established for the maximum allowable trajectory variations. The most restrictive criterion allows a 25% variation in Reynolds number at constant Mach number between trajectories.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 2008-6559 , AIAA Atmospheric Flight Mechanics Conference; Aug 18, 2008 - Aug 21, 2008; Honolulu, HI; United States
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    NASA TECHNICAL REPORTS
  • 5
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    Aerodynamic Shape Optimization of Complex Aircraft Configurations via an Adjoint Formulation (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This work describes the implementation of optimization techniques based on control theory for complex aircraft configurations. Here control theory is employed to derive the adjoint differential equations, the solution of which allows for a drastic reduction in computational costs over previous design methods (13, 12, 43, 38). In our earlier studies (19, 20, 22, 23, 39, 25, 40, 41, 42) it was shown that this method could be used to devise effective optimization procedures for airfoils, wings and wing-bodies subject to either analytic or arbitrary meshes. Design formulations for both potential flows and flows governed by the Euler equations have been demonstrated, showing that such methods can be devised for various governing equations (39, 25). In our most recent works (40, 42) the method was extended to treat wing-body configurations with a large number of mesh points, verifying that significant computational savings can be gained for practical design problems. In this paper the method is extended for the Euler equations to treat complete aircraft configurations via a new multiblock implementation. New elements include a multiblock-multigrid flow solver, a multiblock-multigrid adjoint solver, and a multiblock mesh perturbation scheme. Two design examples are presented in which the new method is used for the wing redesign of a transonic business jet.
    Keywords: Aircraft Design, Testing and Performance
    Type: NASA-CR-203275 , NAS 1.26:203275 , RIACS-TR-96-02 , AIAA Paper 96-0094 , Aerospace Sciences Meeting and Exhibit; Jan 01, 1996; United States
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    NASA TECHNICAL REPORTS
  • 6
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    Insight Aerothermal Environment Assessment (2020)
    In:  CASI
    Details
    Publication Date: 2020-01-17
    Description: The Mars Interior Exploration using Seismic Investigations, Geodesy and Heat Transport (InSight) spacecraft, which successfully touched down on the planet surface on November 26, 2018, was proposed as a near build-to-print copy of the Mars Phoenix vehicle to reduce the overall cost and risk of the mission. Since the lander payload and the atmospheric entry trajectory were similar enough to those of the Phoenix mission, it was expected that the Phoenix thermal protection material thickness would be sufficient to withstand the entry heat load. However, allowances were made for increasing the heatshield thickness because the planned spacecraft arrival date coincided with the Mars dust storm season. The aftbody Thermal Protection System (TPS) components were not expected to change. In a first for a US Mars mission, the aerothermal environments for InSight included estimates of radiative heat flux to the aftbody from the wake. The combined convective and radiative heat fluxes were used to determine if the as-flown Phoenix thermal protection system (TPS) design would be sufficient for InSight. Although the radiative heat fluxes on the aftbody were predicted to be comparable to, or even higher than the local convective heat fluxes, all analyses of the aftbody TPS showed that the design would still be adequate. Aerothermal environments were computed for the vehicle from post-flight reconstruction of the atmosphere and trajectory and compared.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: ARC-E-DAA-TN76667 , AIAA SciTech 2020; Jan 06, 2020 - Jan 10, 2020; Orlando, FL; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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