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  • 1
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    Effects of flight instrumentation errors on the estimation of aircraft stability and control derivatives (1974)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: An error analysis program based on an output error estimation method was used to evaluate the effects of sensor and instrumentation errors on the estimation of aircraft stability and control derivatives. A Monte Carlo analysis was performed using simulated flight data for a high performance military aircraft, a large commercial transport, and a small general aviation aircraft for typical cruise flight conditions. The effects of varying the input sequence and combinations of the sensor and instrumentation errors were investigated. The results indicate that both the parameter accuracy and the corresponding measurement trajectory fit error can be significantly affected. Of the error sources considered, instrumentation lags and control measurement errors were found to be most significant.
    Keywords: AERODYNAMICS
    Type: Parameter Estimation Tech. and Appl. in Aircraft Flight Testing; p 261-280
    Format: text
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  • 2
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    A numerical method for unsteady aerodynamics via acoustics (1991)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Formal solutions to the wave equation may be conveniently described within the framework of generalized function theory. A generalized function theory is used to yield a formulation and formal solution of a wave equation describing oscillation of a flat plate from which a numerical method may be derived.
    Keywords: AERODYNAMICS
    Type: Old Dominion Univ., NASA/American Society for Engineering Ed; Old Dominion Univ.,
    Format: application/pdf
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  • 3
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    Wind tunnel blockage study of a generic three-dimensional sidewall compression scramjet inlet at Mach 10 (1991)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A large scale model of a generic three-dimensional sidewall compression scramjet inlet has been designed based on the results of a computational parametric study for testing in the 31-inch Mach 10 Hypersonic Wind Tunnel at the NASA Langley Research Center. In order to increase the instrumentation density in interaction regions for a highly instrumented model, it is desirable to make the model as large as possible. When the cross-sectional area of a model becomes large relative to the inviscid core size of the tunnel, the effects of blockage must be considered. In order to assess these effects, a blockage model (an inexpensive, much less densely instrumented version of the configuration) was fabricated for preliminary testing. Since it was desired to determine both the effect of the model on the performance of the wind tunnel and also to determine if the inlet would start, the model possessed a total of 32 static pressure orifices distributed on the forebody plane and sidewalls; seventeen static pressure orifices on the tunnel wall and 3 pitot probes on the model monitored the tunnel performance. This paper presents the design considerations in the development of the wind tunnel model and the blockage aspects of the effects of contraction ratio, cowl location, Reynolds number, and angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-0294
    Format: text
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  • 4
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    Solutions of the Navier-Stokes equations in various flow regimes on fields containing any number of arbitrary bodies using boundary-fitted coordinate systems (1976)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Keywords: AERODYNAMICS
    Type: International Conference on Numerical Methods in Fluid Dynamics; Jun 28, 1976 - Jul 02, 1976; Enschede; Netherlands
    Format: text
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  • 5
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    Sandwich Composite, Syntactic Foam Core Based, Application for Space Structures (2000)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The current Solid Rocket Booster (SRB) launch vehicle has several metal based components that require a Thermal Protective System (TPS) be applied to the exterior surface to ensure its structural integrity and to protect the interior hardware from aerodynamic heating. TPS materials have distinct disadvantages associated with their use. One disadvantage to the application of TPS is that it can act as a debris source to the Space Shuttle Orbiter during flight and it also adds weight to the system without directly contributing any structural strength. One of the specific areas examined under this program was to replace a metal/TPS system with polymer based composites. A polymer matrix based sandwich composite was developed which had both structural and insulative properties to meet the high aerodynamic structural and heating load survival requirements. The SRB Nose Cap was selected as a candidate for this application. The sandwich system being qualified for this application is a carbon/epoxy outer and inner skin with a high strength-low thermal conductivity syntactic foam core.
    Keywords: Composite Materials
    Type: 45th SAMPE Symposium; May 21, 2000 - May 25, 2000; Long Beach, CA; United States
    Format: application/pdf
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  • 6
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    Damage Tolerance of Composites (2007)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Fracture control requirements have been developed to address damage tolerance of composites for manned space flight hardware. The requirements provide the framework for critical and noncritical hardware assessment and testing. The need for damage threat assessments, impact damage protection plans, and nondestructive evaluation are also addressed. Hardware intended to be damage tolerant have extensive coupon, sub-element, and full-scale testing requirements in-line with the Building Block Approach concept from the MIL-HDBK-17, Department of Defense Composite Materials Handbook.
    Keywords: Composite Materials
    Type: National Space and Missile Materials Symposium; Jun 25, 2007 - Jun 29, 2007; Keystone, CO; United States
    Format: application/pdf
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  • 7
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    A flutter investigation of all-moveable NASP-like wings at hypersonic speeds (1993)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Six alternative all-moving wing configurations applicable to the NASP hypersonic/transatmospheric vehicle have undergone aeroelasticity testing in NASA-Langley's Mach-20-capable Helium Tunnel that yielded data for such parametric variations as airfoil profile and wing planform, wing-pivot flexure stiffness, and mass imbalance. While all wings fluttered at dynamic pressures lower than predicted by second-order piston-theory aerodynamics, this was of limited amplitude, suggesting nonlinear external-flow behavior. Slab airfoils were more stable than diamond-shaped ones; blunt leading edges enhance stability relative to sharp ones, and stiffer pivolts extert a stabilizing influence.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 93-1315 , AIAA, ASME, ASCE, AHS, and ASC, Structures, Structural Dynamics and Materials Conference, 34th and AIAA and ASME, Adaptive Structures Forum; Apr 19, 1993 - Apr 22, 1993; La Jolla, CA; United States|; 10 p.
    Format: text
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  • 8
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    Evaluation of Microcracking in Two Carbon-Fiber/Epoxy-Matrix Composite Cryogenic Tanks (2001)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Two graphite/epoxy cryogenic pressure vessels were evaluated for microcracking. The X-33 LH2 tank lobe skins were extensively examined for microcracks. Specimens were removed from the inner skin of the X-33 tank for tensile testing. The data obtained from these tests were used to model expected microcrack density as a function of stress. Additionally, the laminate used in the Marshall Space Flight Center (MSFC) Composite Conformal, Cryogenic, Common Bulkhead, Aerogel-Insulated Tank (CBAT) was evaluated. Testing was performed in an attempt to predict potential microcracking during testing of the CBAT.
    Keywords: Composite Materials
    Type: NASA/TM-2001-211194 , M-1024 , NAS 1.15:211194
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  • 9
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    The Effects of Foam Thermal Protection System on the Damage Tolerance Characteristics of Composite Sandwich Structures for Launch Vehicles (2011)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: For any structure composed of laminated composite materials, impact damage is one of the greatest risks and therefore most widely tested responses. Typically, impact damage testing and analysis assumes that a solid object comes into contact with the bare surface of the laminate (the outer ply). However, most launch vehicle structures will have a thermal protection system (TPS) covering the structure for the majority of its life. Thus, the impact response of the material with the TPS covering is the impact scenario of interest. In this study, laminates representative of the composite interstage structure for the Ares I launch vehicle were impact tested with and without the planned TPS covering, which consists of polyurethane foam. Response variables examined include maximum load of impact, damage size as detected by nondestructive evaluation techniques, and damage morphology and compression after impact strength. Results show that there is little difference between TPS covered and bare specimens, except the residual strength data is higher for TPS covered specimens.
    Keywords: Composite Materials
    Type: NASA/TP-2011-216457 , M-1306
    Format: application/pdf
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  • 10
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    Simplification of Fatigue Test Requirements for Damage Tolerance of Composite Interstage Launch Vehicle Hardware (2010)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The issue of fatigue loading of structures composed of composite materials is considered in a requirements document that is currently in place for manned launch vehicles. By taking into account the short life of these parts, coupled with design considerations, it is demonstrated that the necessary coupon level fatigue data collapse to a static case. Data from a literature review of past studies that examined compressive fatigue loading after impact and data generated from this experimental study are presented to support this finding. Damage growth, in the form of infrared thermography, was difficult to detect due to rapid degradation of compressive properties once damage growth initiated. Unrealistically high fatigue amplitudes were needed to fail 5 of 15 specimens before 10,000 cycles were reached. Since a typical vehicle structure, such as the Ares I interstage, only experiences a few cycles near limit load, it is concluded that static compression after impact (CAI) strength data will suffice for most launch vehicle structures.
    Keywords: Composite Materials
    Type: NASA/TP-2010-216434 , M-1283
    Format: application/pdf
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