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Ko, William L.

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

At least 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS2010

Methods for In-Flight Wing Shape Predictions of Highly Flexible Unmanned Aerial Vehicles: Formulation of Ko Displacement Theory

The Ko displacement theory is formulated for a cantilever tubular wing spar under bending, torsion, and combined bending and torsion loading. The Ko displacement equations are expressed in terms of strains measured at multiple sensing stations equally spaced on the surface of the wing spar. The bending and distortion strain data can then be input to the displacement equations to calculate slopes, deflections, and cross-sectional twist angles of the wing spar at the strain-sensing stations for generating the deformed shapes of flexible aircraft wing spars. The displacement equations have been successfully validated for accuracy by finite-element analysis. The Ko displacement theory that has been formulated could also be applied to calculate the deformed shape of simple and tapered beams, plates, and tapered cantilever wing boxes. The Ko displacement theory and associated strain-sensing system (such as fiber optic sensors) form a powerful tool for in-flight deformation monitoring of flexible wings and tails, such as those often employed on unmanned aerial vehicles. Ultimately, the calculated displacement data can be visually displayed in real time to the ground-based pilot for monitoring the deformed shape of unmanned aerial vehicles during flight.

Ko, William L.↗
NASA NTRS2009

Applications of Ko Displacement Theory to the Deformed Shape Predictions of the Doubly-Tapered Ikhana Wing

The Ko displacement theory, formulated for weak nonuniform (slowly changing cross sections) cantilever beams, was applied to the deformed shape analysis of the doubly-tapered wings of the Ikhana unmanned aircraft. The two-line strain-sensing system (along the wingspan) was used for sensing the bending strains needed for the wing-deformed shapes (deflections and cross-sectional twist) analysis. The deflection equation for each strain-sensing line was expressed in terms of the bending strains evaluated at multiple numbers of strain-sensing stations equally spaced along the strain-sensing line. For the preflight shape analysis of the Ikhana wing, the strain data needed for input to the displacement equations for the shape analysis were obtained from the nodal-stress output of the finite-element analysis. The wing deflections and cross-sectional twist angles calculated from the displacement equations were then compared with those computed from the finite-element computer program. The Ko displacement theory formulated for weak nonlinear cantilever beams was found to be highly accurate in the deformed shape predictions of the doubly-tapered Ikhana wing.

Ko, William L.↗
NASA NTRS2009

Further Development of Ko Displacement Theory for Deformed Shape Predictions of Nonuniform Aerospace Structures

The Ko displacement theory previously formulated for deformed shape predictions of nonuniform beam structures is further developed mathematically. The further-developed displacement equations are expressed explicitly in terms of geometrical parameters of the beam and bending strains at equally spaced strain-sensing stations along the multiplexed fiber-optic sensor line installed on the bottom surface of the beam. The bending strain data can then be input into the displacement equations for calculations of local slopes, deflections, and cross-sectional twist angles for generating the overall deformed shapes of the nonuniform beam. The further-developed displacement theory can also be applied to the deformed shape predictions of nonuniform two-point supported beams, nonuniform panels, nonuniform aircraft wings and fuselages, and so forth. The high degree of accuracy of the further-developed displacement theory for nonuniform beams is validated by finite-element analysis of various nonuniform beam structures. Such structures include tapered tubular beams, depth-tapered unswept and swept wing boxes, width-tapered wing boxes, and double-tapered wing boxes, all under combined bending and torsional loads. The Ko displacement theory, combined with the fiber-optic strain-sensing system, provide a powerful tool for in-flight deformed shape monitoring of unmanned aerospace vehicles by ground-based pilots to maintain safe flights.

Ko, William L.↗
NASA NTRS2009

Method for Real-Time Structure Shape-Sensing

The invention is a method for obtaining the displacement of a flexible structure by using strain measurements obtained by stain sensor,. By obtaining the displacement of structures in this manner, one may construct the deformed shape of the structure and display said deformed shape in real-time, enabling active control of the structure shape if desired.

Ko, William L.↗
NASA NTRS2008

Real-time In-Flight Strain and Deflection Monitoring with Fiber Optic Sensors

This viewgraph presentation reviews Dryden's efforts to develop in-flight monitoring based on Fiber Optics. One of the motivating factors for this development was the breakup of the Helios aircraft. On Ikhana the use of fiber optics for wing shape sensing is being developed. They are being used to flight validate fiber optic sensor measurements and real-time wing shape sensing predictions on NASA's Ikhana vehicle; validate fiber optic mathematical models and design tools; Assess technical viability and, if applicable, develop methodology and approach to incorporate wing shape measurements within the vehicle flight control system, and develop and flight validate advanced approaches to perform active wing shape control.

Richards, Lance↗
NASA NTRS2008

Fiber Optic Wing Shape Sensing on NASA's Ikhana UAV

This document discusses the development of fiber optic wing shape sensing on NASA's Ikhana vehicle. The Dryden Flight Research Center's Aerostructures Branch initiated fiber-optic instrumentation development efforts in the mid-1990s. Motivated by a failure to control wing dihedral resulting in a mishap with the Helios aircraft, new wing displacement techniques were developed. Research objectives for Ikhana included validating fiber optic sensor measurements and real-time wing shape sensing predictions; the validation of fiber optic mathematical models and design tools; assessing technical viability and, if applicable, developing methodology and approaches to incorporate wing shape measurements within the vehicle flight control system; and, developing and flight validating approaches to perform active wing shape control using conventional control surfaces and active material concepts.

Richards, Lance↗
NASA NTRS2007

Lunar Return Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Wall Structures

Thermostructural analysis was performed on generic crew exploration vehicle (GCEV) heat shielded wall structures subjected to reentry heating rates based on five potential lunar return reentry trajectories. The GCEV windward outer wall is fabricated with a graphite/epoxy composite honeycomb sandwich panel and the inner wall with an aluminum honeycomb sandwich panel. The outer wall is protected with an ablative Avcoat-5026-39H/CG thermal protection system (TPS). A virtual ablation method (a graphical approximation) developed earlier was further extended, and was used to estimate the ablation periods, ablation heat loads, and the TPS recession layer depths. It was found that up to 83 95 percent of the total reentry heat load was dissipated in the TPS ablation process, leaving a small amount (3-15 percent) of the remaining total reentry heat load to heat the virgin TPS and maintain the TPS surface at the ablation temperature, 1,200 F. The GCEV stagnation point TPS recession layer depths were estimated to be in the range of 0.280-0.910 in, and the allowable minimum stagnation point TPS thicknesses that could maintain the substructural composite sandwich wall at the limit temperature of 300 F were found to be in the range of 0.767-1.538 in. Based on results from the present analyses, the lunar return abort ballistic reentry was found to be quite attractive because it required less TPS weight than the lunar return direct, the lunar return skipping, or the low Earth orbit guided reentry, and only 11.6 percent more TPS weight than the low Earth orbit ballistic reentry that will encounter a considerable weight penalty to obtain the Earth orbit. The analysis also showed that the TPS weight required for the lunar return skipping reentry was much more than the TPS weight necessary for any of the other reentry trajectories considered.

Ko, William L.↗
NASA NTRS2007

Displacement Theories for In-Flight Deformed Shape Predictions of Aerospace Structures

Displacement theories are developed for a variety of structures with the goal of providing real-time shape predictions for aerospace vehicles during flight. These theories are initially developed for a cantilever beam to predict the deformed shapes of the Helios flying wing. The main structural configuration of the Helios wing is a cantilever wing tubular spar subjected to bending, torsion, and combined bending and torsion loading. The displacement equations that are formulated are expressed in terms of strains measured at multiple sensing stations equally spaced on the surface of the wing spar. Displacement theories for other structures, such as tapered cantilever beams, two-point supported beams, wing boxes, and plates also are developed. The accuracy of the displacement theories is successfully validated by finite-element analysis and classical beam theory using input-strains generated by finite-element analysis. The displacement equations and associated strain-sensing system (such as fiber optic sensors) create a powerful means for in-flight deformation monitoring of aerospace structures. This method serves multiple purposes for structural shape sensing, loads monitoring, and structural health monitoring. Ultimately, the calculated displacement data can be visually displayed to the ground-based pilot or used as input to the control system to actively control the shape of structures during flight.

Ko, William L.↗
NASA NTRS2007

Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Structure

Comparative studies were performed on the heat-shielding characteristics of honeycomb-core sandwich panels fabricated with different materials for possible use as wall panels for the proposed crew exploration vehicle. Graphite/epoxy sandwich panel was found to outperform aluminum sandwich panel under the same geometry due to superior heat-shielding qualities and lower material density. Also, representative reentry heat-transfer analysis was performed on the windward wall structures of a generic crew exploration vehicle. The Apollo low Earth orbit reentry trajectory was used to calculate the reentry heating rates. The generic crew exploration vehicle has a graphite/epoxy composite honeycomb sandwich exterior wall and an aluminum honeycomb sandwich interior wall, and is protected with the Apollo thermal protection system ablative material. In the thermal analysis computer program used, the TPS ablation effect was not yet included; however, the results from the nonablation heat-transfer analyses were used to develop a "virtual ablation" method to estimate the ablation heat loads and the thermal protection system recession thicknesses. Depending on the severity of the heating-rate time history, the virtual ablation period was found to last for 87 to 107 seconds and the ablation heat load was estimated to be in the range of 86 to 88 percent of the total heat load for the ablation time period. The thermal protection system recession thickness was estimated to be in the range of 0.08 to 0.11 inches. For the crew exploration vehicle zero-tilt and 18-degree-tilt stagnation points, thermal protection system thicknesses of h = {0.717, 0.733} inches were found to be adequate to keep the substructural composite sandwich temperature below the limit of 300 F.

Ko, William L.↗
NASA NTRS2007

Incorporation of Half-Cycle Theory Into Ko Aging Theory for Aerostructural Flight-Life Predictions

The half-cycle crack growth theory was incorporated into the Ko closed-form aging theory to improve accuracy in the predictions of operational flight life of failure-critical aerostructural components. A new crack growth computer program was written for reading the maximum and minimum loads of each half-cycle from the random loading spectra for crack growth calculations and generation of in-flight crack growth curves. The unified theories were then applied to calculate the number of flights (operational life) permitted for B-52B pylon hooks and Pegasus adapter pylon hooks to carry the Hyper-X launching vehicle that air launches the X-43 Hyper-X research vehicle. A crack growth curve for each hook was generated for visual observation of the crack growth behavior during the entire air-launching or captive flight. It was found that taxiing and the takeoff run induced a major portion of the total crack growth per flight. The operational life theory presented can be applied to estimate the service life of any failure-critical structural components.

Ko, William L.↗
NASA NTRS2006

Extended Aging Theories for Predictions of Safe Operational Life of Critical Airborne Structural Components

The previously developed Ko closed-form aging theory has been reformulated into a more compact mathematical form for easier application. A new equivalent loading theory and empirical loading theories have also been developed and incorporated into the revised Ko aging theory for the prediction of a safe operational life of airborne failure-critical structural components. The new set of aging and loading theories were applied to predict the safe number of flights for the B-52B aircraft to carry a launch vehicle, the structural life of critical components consumed by load excursion to proof load value, and the ground-sitting life of B-52B pylon failure-critical structural components. A special life prediction method was developed for the preflight predictions of operational life of failure-critical structural components of the B-52H pylon system, for which no flight data are available.

Ko, William L.↗
NASA NTRS2005

Stress Analysis of B-52B and B-52H Air-Launching Systems Failure-Critical Structural Components

The operational life analysis of any airborne failure-critical structural component requires the stress-load equation, which relates the applied load to the maximum tangential tensile stress at the critical stress point. The failure-critical structural components identified are the B-52B Pegasus pylon adapter shackles, B-52B Pegasus pylon hooks, B-52H airplane pylon hooks, B-52H airplane front fittings, B-52H airplane rear pylon fitting, and the B-52H airplane pylon lower sway brace. Finite-element stress analysis was performed on the said structural components, and the critical stress point was located and the stress-load equation was established for each failure-critical structural component. The ultimate load, yield load, and proof load needed for operational life analysis were established for each failure-critical structural component.

Ko, William L.↗
NASA NTRS2004

Hypothetical Reentry Thermostructural Performance of Space Shuttle Orbiter With Missing or Eroded Thermal Protection Tiles

This report deals with hypothetical reentry thermostructural performance of the Space Shuttle orbiter with missing or eroded thermal protection system (TPS) tiles. The original STS-5 heating (normal transition at 1100 sec) and the modified STS-5 heating (premature transition at 800 sec) were used as reentry heat inputs. The TPS missing or eroded site is assumed to be located at the center or corner (spar-rib juncture) of the lower surface of wing midspan bay 3. For cases of missing TPS tiles, under the original STS-5 heating, the orbiter can afford to lose only one TPS tile at the center or two TPS tiles at the corner (spar-rib juncture) of the lower surface of wing midspan bay 3. Under modified STS-5 heating, the orbiter cannot afford to lose even one TPS tile at the center or at the corner of the lower surface of wing midspan bay 3. For cases of eroded TPS tiles, the aluminum skin temperature rises relatively slowly with the decreasing thickness of the eroded central or corner TPS tile until most of the TPS tile is eroded away, and then increases exponentially toward the missing tile case.

Ko, William L.↗
NASA NTRS2004

Debonding Stress Concentrations in a Pressurized Lobed Sandwich-Walled Generic Cryogenic Tank

A finite-element stress analysis has been conducted on a lobed composite sandwich tank subjected to internal pressure and cryogenic cooling. The lobed geometry consists of two obtuse circular walls joined together with a common flat wall. Under internal pressure and cryogenic cooling, this type of lobed tank wall will experience open-mode (a process in which the honeycomb is stretched in the depth direction) and shear stress concentrations at the junctures where curved wall changes into flat wall (known as a curve-flat juncture). Open-mode and shear stress concentrations occur in the honeycomb core at the curve-flat junctures and could cause debonding failure. The levels of contributions from internal pressure and temperature loading to the open-mode and shear debonding failure are compared. The lobed fuel tank with honeycomb sandwich walls has been found to be a structurally unsound geometry because of very low debonding failure strengths. The debonding failure problem could be eliminated if the honeycomb core at the curve-flat juncture is replaced with a solid core.

Ko, William L.↗
NASA NTRS2004

Heat Shielding Characteristics and Thermostructural Performance of a Superalloy Honeycomb Sandwich Thermal Protection System (TPS)

Heat-transfer, thermal bending, and mechanical buckling analyses have been performed on a superalloy "honeycomb" thermal protection system (TPS) for future hypersonic flight vehicles. The studies focus on the effect of honeycomb cell geometry on the TPS heat-shielding performance, honeycomb cell wall buckling characteristics, and the effect of boundary conditions on the TPS thermal bending behavior. The results of the study show that the heat-shielding performance of a TPS panel is very sensitive to change in honeycomb core depth, but insensitive to change in honeycomb cell cross-sectional shape. The thermal deformations and thermal stresses in the TPS panel are found to be very sensitive to the edge support conditions. Slight corrugation of the honeycomb cell walls can greatly increase their buckling strength.

Ko, William L.↗
NASA NTRS2003

Aging Theories for Establishing Safe Life Spans of Airborne Critical Structural Components

New aging theories have been developed to establish the safe life span of airborne critical structural components such as B-52B aircraft pylon hooks for carrying air-launch drop-test vehicles. The new aging theories use the equivalent-constant-amplitude loading spectrum to represent the actual random loading spectrum with the same damaging effect. The crack growth due to random loading cycling of the first flight is calculated using the half-cycle theory, and then extrapolated to all the crack growths of the subsequent flights. The predictions of the new aging theories (finite difference aging theory and closed-form aging theory) are compared with the classical flight-test life theory and the previously developed Ko first- and Ko second-order aging theories. The new aging theories predict the number of safe flights as considerably lower than that predicted by the classical aging theory, and slightly lower than those predicted by the Ko first- and Ko second-order aging theories due to the inclusion of all the higher order terms.

Ko, William L.↗
NASA NTRS2001

Thermostructural Analysis of Unconventional Wing Structures of a Hyper-X Hypersonic Flight Research Vehicle for the Mach 7 Mission

Heat transfer, thermal stresses, and thermal buckling analyses were performed on the unconventional wing structures of a Hyper-X hypersonic flight research vehicle (designated as X-43) subjected to nominal Mach 7 aerodynamic heating. A wing midspan cross section was selected for the heat transfer and thermal stress analyses. Thermal buckling analysis was performed on three regions of the wing skin (lower or upper); 1) a fore wing panel, 2) an aft wing panel, and 3) a unit panel at the middle of the aft wing panel. A fourth thermal buckling analysis was performed on a midspan wing segment. The unit panel region is identified as the potential thermal buckling initiation zone. Therefore, thermal buckling analysis of the Hyper-X wing panels could be reduced to the thermal buckling analysis of that unit panel. "Buckling temperature magnification factors" were established. Structural temperature-time histories are presented. The results show that the concerns of shear failure at wing and spar welded sites, and of thermal buckling of Hyper-X wing panels, may not arise under Mach 7 conditions.

Ko, William L.↗