Engineering topics

Ko, W. L.

Publications and source records attributed to Ko, W. L..

At least 19 records

NASA NTRS1993

Open-Mode Delamination In Horseshoelike Specimens

Tests to determine open-mode-delamination stresses of laminated composite materials improved by use of horseshoelike curved bar specimens. Two horseshoelike specimen shapes proposed: one with semicircular midsection and straight end sections, and one essentially half of ellipse cut along its minor axis. Shapes suitable for both solid and sandwich laminates. Calculations show elliptical horseshoe bar concentrates delamination stress slightly more than semicircular/straight bar does. Semicircular/straight bar specimen used to determine both delamination strength of solid laminate and debonding strength of sandwich laminate.

Ko, W. L.↗

NASA NTRS1989

Application of fracture mechanics and half-cycle method to the prediction of fatigue life of B-52 aircraft pylon components

Stress intensity levels at various parts of the NASA B-52 carrier aircraft pylon were examined for the case when the pylon store was the space shuttle solid rocket booster drop test vehicle. Eight critical stress points were selected for the pylon fatigue analysis. Using fracture mechanics and the half-cycle theory (directly or indirectly) for the calculations of fatigue-crack growth ,the remaining fatigue life (number of flights left) was estimated for each critical part. It was found that the two rear hooks had relatively short fatigue life and that the front hook had the shortest fatigue life of all the parts analyzed. The rest of the pylon parts were found to be noncritical because of their extremely long fatigue life associated with the low operational stress levels.

Ko, W. L.↗

NASA NTRS1986

Buckling behavior of Rene 41 tubular panels for a hypersonic aircraft wing

The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panel. In spite of some data scattering, resulting from large extrapolations of the data fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.

Ko, W. L.↗

NASA NTRS1986

Buckling behavior of Rene 41 tubular panels for a hypersonic aircraft wing

The buckling characteristics of Rene 41 tubular panels for a hypersonic aircraft wing were investigated. The panels were repeatedly tested for buckling characteristics using a hypersonic wing test structure and a universal tension/compression testing machine. The nondestructive buckling tests were carried out under different combined load conditions and in different temperature environments. The force/stiffness technique was used to determine the buckling loads of the panels. In spite of some data scattering resulting from large extrapolations of the data-fitting curve (because of the termination of applied loads at relatively low percentages of the buckling loads), the overall test data correlate fairly well with theoretically predicted buckling interaction curves. Also, the structural efficiency of the tubular panels was found to be slightly higher than that of beaded panels.

Ko, W. L.↗

NASA NTRS1985

Stress concentration around a small circular hole in the HiMAT composite plate

Anisotropic plate theory is used to calculate the anisotropic stress concentration factors for a composite plate (AS/3501-5 graphite/epoxy composite, single ply or laminated) containing a circular hole. This composite material is used on the highly maneuverable aircraft technology (HiMAT) vehicle. It is found that the anisotropic stress concentration factor could be greater or less than 3 (the stress concentration factor for isotropic materials), and that the locations of the maximum tangential stress points could shift with the change of fiber orientation with respect to the loading axis. The effect of hole size on the stress concentration factor is examined using the Point Stress Criterion and the Averaged Stress Criterion. The predicted stress concentration factors based on the two theories compared fairly well with the measured values for the hole size 0.3175 cm (1/8 in). It is also found that through the lamination process, the stress concentration factor could be reduced drastically, indicating an improvement in structural performance.

Ko, W. L.↗

NASA NTRS1985

Stress analyses of B-52 pylon hooks

The NASTRAN finite element computer program was used in the two dimensional stress analysis of B-52 carrier aircraft pylon hooks: (1) old rear hook (which failed), (2) new rear hook (improved geometry), (3) new DAST rear hook (derated geometry), and (4) front hook. NASTRAN model meshes were generated by the aid of PATRAN-G computer program. Brittle limit loads for all the four hooks were established. The critical stress level calculated from NASTRAN agrees reasonably well with the values predicted from the fracture mechanics for the failed old rear hook.

Ko, W. L.↗

NASA NTRS1985

Impacts of Space Shuttle thermal protection system tile on F-15 aircraft vertical tile

Impacts of the space shuttle thermal protection system (TPS) tile on the leading edge and the side of the vertical tail of the F-15 aircraft were analyzed under different TPS tile orientations. The TPS tile-breaking tests were conducted to simulate the TPS tile impacts. It was found that the predicted tile impact forces compare fairly well with the tile-breaking forces, and the impact forces exerted on the F-15 aircraft vertical tail were relatively low because a very small fraction of the tile kinetic energy was dissipated in the impact, penetration, and fracture of the tile. It was also found that the oblique impact of the tile on the side of the F-15 aircraft vertical tail was unlikely to dent the tail surface.

Ko, W. L.↗

NASA NTRS1984

Thermal response of Space Shuttle wing during reentry heating

A structural performance and resizing (SPAR) finite element thermal analysis computer program was used in the heat transfer analysis of the Space Shuttle Orbiter that was subjected to reentry aerodynamic heatings. One wing segment of the right wing (WS 240) and the whole left wing were selected for the thermal analysis. Results showed that the predicted thermal protection system (TPS) temperatures were in good agreement with the space transportation system, trajectory 5 (STS-5) flight-measured temperatures. In addition, calculated aluminum structural temperatures were in fairly good agreement with flight data up to the point of touchdown. Results also showed that the internal free convection has a considerable effect on the change of structural temperatures after touchdown.

Gong, L.↗

NASA NTRS1984

Preflight transient dynamic analyses of B-52 aircraft carrying Space Shuttle solid rocket booster drop-test vehicle

This paper concerns the transient dynamic analysis of the B-52 aircraft carrying the Space Shuttle solid rocket booster drop test vehicle (SRB/DTV). The NASA structural analysis (NASTRAN) finite element computer program was used in the analysis. The B-52 operating conditions considered for analysis were (1) landing and (2) braking on aborted takeoff runs. The transient loads for the B-52 pylon front and rear hooks were calculated. The results can be used to establish the safe maneuver envelopes for the B-52 carrying the SRB/DTV in landings and brakings.

Ko, W. L.↗

NASA NTRS1984

Thermal response of Space Shuttle wing during reentry heating

A structural performance and resizing (SPAR) finite element thermal analysis computer program was used in the heat transfer analysis of the space shuttle orbiter that was subjected to reentry aerodynamic heatings. One wing segment of the right wing (WS 240) and the whole left wing were selected for the thermal analysis. Results showed that the predicted thermal protection system (TPS) temperatures were in good agreement with the space transportation system, trajectory 5 (STS-5) flight-measured temperatures. In addition, calculated aluminum structural temperatures were in fairly good agreement with the flight data up to the point of touchdown. Results also showed that the internal free convection had a considerable effect on the change of structural temperatures after touchdown.

Gong, L.↗

NASA NTRS1983

Pre-flight transient dynamic analysis of B-52 carrying Space Shuttle solid rocket booster drop-test vehicle

This paper concerns the transient dynamic analysis of the B-52 aircraft carrying the Space Shuttle solid-rocket booster drop-test vehicle (SRB/DTV). The NASA structural analysis (NASTRAN) finite-element computer program was used in the analysis. The B-52 operating conditions considered for analysis were (1) landing and (2) braking on aborted takeoff runs. The transient loads for the B-52 pylon front and rear hooks were calculated. The results can be used to establish the safe maneuver envelopes for the B-52 carrying the SRB/DTV in landings and brakings.

Ko, W. L.↗

NASA NTRS1982

Metal Sandwith Panel With Biaxially Corrugated Core

Biaxially-corrugated sandwich core, formed by concurrent diffusion bonding and superplastic deformation, makes a proposed sandwich panel unusually strong and stiff. New panel should be useful where light-weight panel that resists bending is needed. Panel would have internal core consisting of multitude of hollow, truncated pyramids. Four layers of panel would be diffusion-binded at all contacting areas.

Ko, W. L.↗

NASA NTRS1982

Reentry heating analysis of space shuttle with comparison of flight data

Surface heating rates and surface temperatures for a space shuttle reentry profile were calculated for two wing cross sections and one fuselage cross section. Heating rates and temperatures at 12 locations on the wing and 6 locations on the fuselage are presented. The heating on the lower wing was most severe, with peak temperatures reaching values of 1240 C for turbulent flow and 900 C for laminar flow. For the fuselage, the most severe heating occured on the lower glove surface where peak temperatures of 910 C and 700 C were calculated for turbulent flow and laminar flow, respectively. Aluminum structural temperatures were calculated using a finite difference thermal analyzer computer program, and the predicted temperatures are compared to measured flight data. Skin temperatures measured on the lower surface of the wing and bay 1 of the upper surface of the wing agreed best with temperatures calculated assuming laminar flow. The measured temperatures at bays two and four on the upper surface of the wing were in quite good agreement with the temperatures calculated assuming separated flow. The measured temperatures on the lower forward spar cap of bay four were in good agreement with values predicted assuming laminar flow.

Gong, L.↗

NASA NTRS1982

Reentry heat transfer analysis of the space shuttle orbiter

A structural performance and resizing finite element thermal analysis computer program was used in the reentry heat transfer analysis of the space shuttle. Two typical wing cross sections and a midfuselage cross section were selected for the analysis. The surface heat inputs to the thermal models were obtained from aerodynamic heating analyses, which assumed a purely turbulent boundary layer, a purely laminar boundary layer, separated flow, and transition from laminar to turbulent flow. The effect of internal radiation was found to be quite significant. With the effect of the internal radiation considered, the wing lower skin temperature became about 39 C (70 F) lower. The results were compared with fight data for space transportation system, trajectory 1. The calculated and measured temperatures compared well for the wing if laminar flow was assumed for the lower surface and bay one upper surface and if separated flow was assumed for the upper surfaces of bays other than bay one. For the fuselage, good agreement between the calculated and measured data was obtained if laminar flow was assumed for the bottom surface. The structural temperatures were found to reach their peak values shortly before touchdown. In addition, the finite element solutions were compared with those obtained from the conventional finite difference solutions.

Ko, W. L.↗

NASA NTRS1982

Radiation from an open-ended waveguide with beam equalizer - A spectral domain analysis

A septum and an impedance matching post are used as a beam equalizer in an open-ended waveguide feed for reflectors used in satellite communications systems. The performance of this design over a frequency band is evaluated using a spectral domain approach. The computed radiation patterns in the E- and H-planes, as well as the results for the impedance match are presented.

Ko, W. L.↗

NASA NTRS1981

Preflight reentry heat transfer analysis of Space Shuttle

Preflight predictions of the structural temperature distributions during entry are compared with data from the initial Shuttle flight. Finite element thermal analysis programming was used to model the heat flow on Shuttle structures and actual gas properties of air were employed in the analyses of aerodynamic heating. Laminar, separated, and turbulent heat fluxes were calculated for varying locations on the craft using velocity-attitude and angle-of-attack projections taken from the nominal STS-1 trajectory. Temperature time histories of the first flight are compared with laminar and turbulent flow assumptions and an unpredicted rapid cooling 1800 sec into entry is credited to inaccurate assumptions of structural heat dissipative properties or flow conditions in that time phase of the flight; additional discrepancies in descriptions of heating of the upper fuselage are attributed to a lack of knowledge of the complex flow patterns existing over that area of the Shuttle body.

Ko, W. L.↗

NASA NTRS1981

Superplastically formed diffusion bonded metallic structure

A metallic sandwich structure particularly suited for use in aerospace industries comprising a base plate, a cover plate, and an orthogonally corrugated core is described. A pair of core plates formed of a superplastic alloy are interposed between the base plate and the cover plate and bonded. Each of the core plates is characterized by a plurality of protrusions comprising square-based, truncated pyramids uniformly aligned along orthogonally related axes perpendicularly bisecting the legs of the bases of the pyramids and alternately inverted along orthogonally related planes diagonally bisecting the pyramids, whereby an orthogonally corrugated core is provided.

Ko, W. L.↗

NASA NTRS1980

Comparison of structural behavior of superplastically formed/diffusion-bonded sandwich structures and honeycomb core sandwich structures

A superplasticity formed/diffusion-bonded (SPF/DB) orthogonally corrugated core sandwich structure is discussed and its structural behavior is compared to that of a conventional honeycomb core sandwich structure. The stiffness and buckling characteristics of the two types of sandwich structures are compared under conditions of equal structural density. It is shown that under certain conditions, the SPF/DB orthogonally corrugated core sandwich structure is slightly more efficient than the optimum honeycomb core (square-cell core) sandwich structure. However, under different conditions, this effect can be reversed.

Ko, W. L.↗