Search NASA⌕ Search

SEARCH · Search NASA

Results for “Design Load”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Voyager design and flight loads comparison

Estimates of flight loads for Voyager 1 and Voyager 2 are summarized. These member loads are obtained by using the measured flight accelerations at the launch vehicle/spacecraft interface as forcing functions for the Voyager mathematical model. The flight loads are compared to the Voyager design loads obtained from the shock spectra/impedance method and to the loads obtained using space vehicle transient loads analysis. Finally, based on these data, an assessment of the shock spectra/impedance loads method used for Voyager is presented.

Chen, J. C.↗
NASA NTRS1992

Determining design gust loads for nonlinear aircraft similarity between methods based on matched filter theory and on stochastic simulation

This is a work-in-progress paper. It explores the similarity between the results from two different analysis methods - one deterministic, the other stochastic - for computing maximized and time-correlated gust loads for nonlinear aircraft. To date, numerical studies have been performed using two different nonlinear aircraft configurations. These studies demonstrate that results from the deterministic analysis method are realizable in the stochastic analysis method.

Scott, Robert C.↗
NASA NTRS1980

Structural dynamics payload loads estimates

Present analytical techniques by which design loads are predicted are very costly and time consuming. Chapter I presents the standard techniques used to analyze payload/booster systems. They are full scale methods in the sense that they all require the solution of the coupled equations of motion of the payload/booster system. Chapter II identifies several short cut methodologies. These already existing techniques do not require the solution of the coupled system equations. The potentials and shortcomings of each of these methods are discussed. Chapter III covers the favored methods accompanied by recommendations for further development, refinement, and demonstrations. An outline of a new approach is also included.

Engels, R. C.↗
NASA NTRS1989

Dynamic loading of spur gears with linear or parabolic tooth profile modification

A computer simulation was conducted to investigate the effects of both linear and parabolic tooth profile modification on the dynamic response of low-contact-ratio spur gears. The effect of the total amount of modification and the length of the modification zone were studied at various loads and speeds to find the optimal profile modification for minimal dynamic loading. Design charts consisting of normalized maximum dynamic load curves were generated for gear systems operated at various loads and with different tooth profile modification. An optimum profile modification can be determined from these design charts to minimize the dynamic loads of spur gear systems.

Lin, Hsiang Hsi↗
NASA NTRS1989

Dynamic loading of spur gears with linear or parabolic tooth profile modifications

A computer simulation was conducted to investigate the effects of both linear and parabolic tooth profile modification on the dynamic response of low-contact-ratio spur gears. The effect of the total amount of modification and the length of the modification zone were studied at various loads and speeds to find the optimal profile modification for minimal dynamic loading. Design charts consisting of normalized maximum dynamic load curves were generated for gear systems operated at various loads and with different tooth profile modification. An optimum profile modification can be determined from these design charts to minimize the dynamic loads of spur gear systems.

Lin, H. H.↗
NASA NTRS1992

National Launch System cycle 1 loads and models data book

This document contains preliminary cycle 1 loads for the National Launch System (NLS) 1 and 2 vehicles. The loads provided and recommended as design loads represent the maximum load expected during prelaunch and flight regimes, i.e., limit loads, except that propellant tank ullage pressure has not been included. Ullage pressure should be added to the loads book values for cases where the addition results in higher loads. The loads must be multiplied by the appropriate factors of safety to determine the ultimate loads for which the structure must be capable.

Bugg, F.↗
NASA NTRS2018

Norton-Thevenin Receptance Coupling (NTRC) as a Payload Analysis Tool

The NASA Engineering and Safety Center (NESC) has funded a study of a new method formulated by NASA Engineers called Norton-Thevenin Receptance Coupling (NTRC) to perform coupled loads analysis (CLA). The problem that NTRC attempts to solve is the dependency of the payload organization to high CLA costs, long schedules, lack of standard capabilities to evaluate multiple configurations and unavailability of loads when needed. NTRC solves the problem by providing a tool that payload developers can use to obtain loads at a fraction of the cost of a CLA at any time that it is required. While NTRC is not intended to replace the formal load cycles performed by the launch vehicle (LV) provider, it will provide the ability to reduce the conservatism in defining preliminary design loads, assess the impact of design changes between formal load cycles, perform trade studies and perform parametric loads analysis where many different design configurations can be evaluated with a minimum amount of data required from the LV provider. NTRC condenses all the necessary information into the launch vehicle to payload/s connection points or boundary degrees of freedom (BD). The launch vehicle model is represented by its impedance at its BDs; its forcing functions are represented by the acceleration at those BDs when the payload is absent and the latter is represented by its impedance at the same BDs. Payload responses are represented by transfer functions of selected response to interface BDs. The methodology has contributed to the Loads and Dynamics discipline advancement and successfully passed Peer Reviews. NTRC is exact in the frequency domain. Time domain replication and accuracy is outstanding. A second phase is envisioned to benchmark the whole set of CLA events for the Agency's most utilized Launch Vehicles, and ready it for operational deployment at NASA.

Kaufman, Daniel↗
NASA NTRS2000

Engine System Loads Development for the Fastrac 60K Flight Engine

Early implementation of structural dynamics finite element analyses for calculation of design loads is considered common design practice for high volume manufacturing industries such as automotive and aeronautical industries. However, with the rarity of rocket engine development programs starts, these tools are relatively new to the design of rocket engines. In the new Fastrac engine program, the focus has been to reduce the cost to weight ratio; current structural dynamics analysis practices were tailored in order to meet both production and structural design goals. Perturbation of rocket engine design parameters resulted in a number of Fastrac load cycles necessary to characterize the impact due to mass and stiffness changes. Evolution of loads and load extraction methodologies, parametric considerations and a discussion of load path sensitivities are discussed.

Frady, Greg↗
NASA NTRS1978

Load and dynamic assessment of B-52B-008 carrier aircraft for finned configuration 1 space shuttle solid rocket booster decelerator subsystem drop test vehicle. Volume 3: Pylon load data method 1

The pylon loading at the drop test vehicle and wing interface attach points is presented. The loads shown are determined using a stiffness method, which assumes the side stiffness of the foreward hook guide to be one-fourth of the fore and aft stiffness of each drag pin. The net effect of this assumption is that the forward hook guide reacts approximately 85% of the drop test vehicle yawing moment. For a comparison of these loads to previous X-15 analysis design loadings, see Volume 1 of this document.

Quade, D. A.↗
NASA NTRS1978

Load and dynamic assessment of B-52B-008 carrier aircraft for finned configuration 1 space shuttle solid rocket booster deceleration subsystem drop test vehicle. Volume 4: Pylon load data

The pylon loading at the drop test vehicle and wing interface attack points is presented. The loads shown are determined using a stiffness method, which assumes the side stiffness of the forward hook guide and the fore and aft stiffness of each drag pin to be equal. The net effect of this assumption is that the forward hook guide reacts approximately 96% of the drop test vehicle yawing moment. For a comparison of these loads to previous X-15 analysis design loadings, see Volume 1 of this document.

Quade, D. A.↗
NASA NTRS2014

Hybrid-Wing-Body Vehicle Composite Fuselage Analysis and Case Study

Recent progress in the structural analysis of a Hybrid Wing-Body (HWB) fuselage concept is presented with the objective of structural weight reduction under a set of critical design loads. This pressurized efficient HWB fuselage design is presently being investigated by the NASA Environmentally Responsible Aviation (ERA) project in collaboration with the Boeing Company, Huntington Beach. The Pultruded Rod-Stiffened Efficient Unitized Structure (PRSEUS) composite concept, developed at the Boeing Company, is approximately modeled for an analytical study and finite element analysis. Stiffened plate linear theories are employed for a parametric case study. Maximum deflection and stress levels are obtained with appropriate assumptions for a set of feasible stiffened panel configurations. An analytical parametric case study is presented to examine the effects of discrete stiffener spacing and skin thickness on structural weight, deflection and stress. A finite-element model (FEM) of an integrated fuselage section with bulkhead is developed for an independent assessment. Stress analysis and scenario based case studies are conducted for design improvement. The FEM model specific weight of the improved fuselage concept is computed and compared to previous studies, in order to assess the relative weight/strength advantages of this advanced composite airframe technology

Mukhopadhyay, Vivek↗
NASA NTRS1993

Improved estimation of random vibration loads in launch vehicles

Random vibration induced load is an important component of the total design load environment for payload and launch vehicle components and their support structures. The current approach to random vibration load estimation is based, particularly at the preliminary design stage, on the use of Miles' equation which assumes a single degree-of-freedom (DOF) system and white noise excitation. This paper examines the implications of the use of multi-DOF system models and response calculation based on numerical integration using the actual excitation spectra for random vibration load estimation. The analytical study presented considers a two-DOF system and brings out the effects of modal mass, damping and frequency ratios on the random vibration load factor. The results indicate that load estimates based on the Miles' equation can be significantly different from the more accurate estimates based on multi-DOF models.

Mehta, R.↗
NASA NTRS2020

InSight Aerothermal Environment Assessment

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

Beck, Robin↗
NASA NTRS2019

InSight's Reconstructed Aerothermal Environments

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

radiation↗
NASA NTRS1975

A shock spectra and impedance method to determine a bound for spacecraft structural loads

This paper proposes a method for the determination of design loads of a spacecraft mounted on a launch vehicle when the significant frequency ranges of the two structures overlap. The derived loads are realistic upper bounds of the flight loads. The method utilizes shock spectra of the spacecraft/launch vehicle interface directional accelerations for each event separately and introduces a reduction factor for the shock spectra which is based on the relative impedance of the spacecraft and launch vehicle. The main advantages of the method are low cost of analysis and timeliness in the design. These advantages are attained with only a slight expense of structural weight compared to the more exact transient analysis method.

Bamford, R.↗