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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.

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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed Simulink model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the Simulink model vs. a full geometric ADAMS model are discussed as well.

Konno, Kevin E.↗
NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed 2-D (Simulink) model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the 2-D model vs. a full 3-D (ADAMS) model are discussed as well.

Konno, Kevin E.↗
NASA NTRS2008

Evaluation of Separation Mechanism Design for the Orion/Ares Launch Vehicle

As a part of the preliminary design work being performed for the Orion vehicle, the Orion to Spacecraft Adaptor (SA) separation mechanism was analyzed and sized, with findings presented here. Sizing is based on worst case abort condition as a result of an anomaly driving the launch vehicle engine thrust vector control hard-over causing a severe vehicle pitch over. This worst case scenario occurs just before Upper Stage Main Engine Cut-Off (MECO) when the vehicle is the lightest and the damping effect due to propellant slosh has been reduced to a minimum. To address this scenario and others, two modeling approaches were invoked. The first approach was a detailed 2-D (Simulink) model to quickly assess the Service Module Engine nozzle to SA clearance for a given separation mechanism. The second approach involved the generation of an Automatic Dynamic Analysis of Mechanical Systems (ADAMS) model to assess secondary effects due to mass centers of gravity that were slightly off the vehicle centerline. It also captured any interference between the Solar Arrays and the Spacecraft Adapter. A comparison of modeling results and accuracy are discussed. Most notably, incorporating a larger SA flange diameter allowed for a natural separation of the Orion and its engine nozzle even at relatively large pitch rates minimizing the kickoff force. Advantages and disadvantages of the 2-D model vs. a full 3-D (ADAMS) model are discussed as well.

Konno, Kevin E.↗
NASA NTRS2026

Mechanical Design and Operation of a Novel Lunar Environment Structural Test Rig (LESTR)

The Lunar Environment Structural Test Rig (LESTR) was developed to address a critical gap in mechanical property data for metal alloy wire materials under lunar-relevant conditions down to 40 K. Conventional aerospace material databases provide thermophysical properties for bulk metals over a wide temperature range, but validated mechanical and physical property data below 77 K, particularly for small-diameter wires, remain limited and are generally the exception rather than the rule. These conditions are essential for Artemis mission hardware such as shape memory alloy (SMA) rover tires. LESTR’s design requirements were to combine a high-stiffness electrodynamic load frame with closed- cycle cryogenic cooling, high-vacuum capability (10–6 torr), and noncontact optical strain measurement to enable tensile, four-point bend, and fatigue testing of wires or other materials and geometries at temperatures from 40 to 125 K. These considerable requirements were merged with the need to lower the barrier of testing for the end user as measured in terms of cost-per-test cycle, safety improvement, and reduction in upkeep costs associated with state-of-the-art solutions associated with cryomechanical material characterization. The system incorporates modular gripping and alignment fixtures; precision thermal management using cryocoolers and embedded heaters; and integrated instrumentation for displacement, load, temperature, and vacuum control. Calibration procedures establish correlations between tooling and specimen temperature, ensuring accurate thermal conditions across the design envelope unbound by cryofluid conditions in heritage immersion-based test systems. Initial validation tests using Inconel (Special Metals Corp.) wire demonstrated accurate ultimate strength and post-yield behavior. The load frame operation was also verified under representative service conditions, including thermal cycling and prolonged fatigue loading. By generating mechanical property data at ultralow temperatures, LESTR fills a critical gap in existing materials databases and provides a scalable platform for iterative alloy development and durability assessment for planetary hardware. This capability supports NASA’s long-term objectives for surface exploration by enabling design confidence for components operating in extreme cryogenic environments.

LESTR↗
NASA NTRS2005

Mechanical Design of Carbon Ion Optics

Carbon Ion Optics are expected to provide much longer thruster life due to their resistance to sputter erosion. There are a number of different forms of carbon that have been used for fabricating ion thruster optics. The mechanical behavior of carbon is much different than that of most metals, and poses unique design challenges. In order to minimize mission risk, the behavior of carbon must be well understood, and components designed within material limitations. Thermal expansion of the thruster structure must be compatible with thermal expansion of the carbon ion optics. Specially designed interfaces may be needed so that grid gap and aperture alignment are not adversely affected by dissimilar material properties within the thruster. The assembled thruster must be robust and tolerant of launch vibration. The following paper lists some of the characteristics of various carbon materials. Several past ion optics designs are discussed, identifying strengths and weaknesses. Electrostatics and material science are not emphasized so much as the mechanical behavior and integration of grid electrodes into an ion thruster.

Haag, Thomas↗
NASA NTRS2018

KaRIn Alignment Mechanism Design, Development, and Testing

The Ka-Band Radar Interferometric System (KaRIn) will characterize the ocean mesoscale/submesoscale circulation and provide a global inventory of all lakes, reservoirs, wetlands, and major rivers. There are two identical Alignment Mechanisms (AM) mirrored on the instrument to correct KaRIn Azimuth radar pointing errors. Each AM is a one-degree-of-freedom precision mechanism. The mechanism must meet challenging thermoelastic distortion requirements over a wide temperature range. The flight design was qualified in less than two years, and the flight hardware is currently in acceptance testing. Discussed are key design details along with supporting development and qualification test results.

Wolff, John Luke↗
NASA NTRS1974

A summary of the mechanical design, testing and performance of the IMP-H and J attitude control systems

The main aspects of the attitude control system used on both the IMP-H and J spacecraft are presented. The mechanical configuration is described. Information on all the specific components comprising the flight system is provided. The acceptance and qualification testing of both individual components and the installed system are summarized. Functional information regarding the operation and performance in relation to the orbiting spacecraft and its mission is included. Related topics which are discussed are: (1) safety requirements, (2) servicing procedures, (3) anomalous behavior, and (4) pyrotechnic devices.

Metzger, J. R.↗
NASA NTRS1995

FORTE antenna element and release mechanism design

The Fast On-Orbit Recording of Transient Events (FORTE) satellite being built by Los Alamos National Laboratory (LANL) and Sandia National Laboratories (SNL) has as its most prominent feature a large deployable (11 m by 5 m) log periodic antenna to monitor emissions from electrical storms on the Earth. This paper describes the antenna and the design for the long elements and explains the dynamics of their deployment and the damping system employed. It also describes the unique paraffin-actuated reusable tie-down and release mechanism employed in the system.

Rohweller, David J.↗
NASA NTRS1992

MSFC Three Point Docking Mechanism design review

In the next few decades, we will be launching expensive satellites and space platforms that will require recovery for economic reasons, because of initial malfunction, servicing, repairs, or out of a concern for post lifetime debris removal. The planned availability of a Three Point Docking Mechanism (TPDM) is a positive step towards an operational satellite retrieval infrastructure. This study effort supports NASA/MSFC engineering work in developing an automated docking capability. The work was performed by the Grumman Space & Electronics Group as a concept evaluation/test for the Tumbling Satellite Retrieval Kit. Simulation of a TPDM capture was performed in Grumman's Large Amplitude Space Simulator (LASS) using mockups of both parts (the mechanism and payload). Similar TPDM simulation activities and more extensive hardware testing was performed at NASA/MSFC in the Flight Robotics Laboratory and Space Station/Space Operations Mechanism Test Bed (6-DOF Facility).

Schaefer, Otto↗
DOE OSTI2024

Mechanical Design Guidelines to Inhibit Fracture in Perovskite Solar Cells

Perovskite (PVSK) solar cells offer significant benefits over conventional silicon cells including low-cost solution processibility, minimal materials usage related to strong photon absorption in thin-film cell architectures, and a tunable bandgap. However, PVSK films are mechanically fragile, and fracture of PVSK layers and adjacent interfaces are a significant concern during fabrication, encapsulation, and operation. Herein, a thin-film mechanics fracture analysis tailored for p–i–n and n–i–p PVSK solar cells on both soda lime glass and polyimide substrates fabricated with three PVSK crystallization methods is presented. Here, the role of thermal processing of each cell layer is explored to determine the maximum allowable temperature below which fracture is inhibited. In the analysis, the mechanics basis for processing and materials selection guidelines for preventing fracture in PVSK solar cells is provided.

adhesion↗
NASA NTRS1992

Mechanical Design Engineering Enabler Project wheel and wheel drives

Our group was assigned the responsibility of designing the wheel and wheel drive system for a proof-of-concept model of the lunar-based ENABLER. ENABLER is a multi-purpose, six wheeled vehicle designed to lift and transport heavy objects associated with the construction of a lunar base. The resulting design was based on the performance criteria of the ENABLER. The drive system was designed to enable the vehicle to achieve a speed of 7 mph on a level surface, climb a 30 percent grade, and surpass a one meter high object and one meter wide crevice. The wheel assemblies were designed to support the entire weight of the vehicle on two wheels. The wheels were designed to serve as the main component of the vehicle's suspension and will provide suitable traction for lunar-type surfaces. The expected performance of the drive system for the ENABLER was influenced by many mechanical factors. The expected top speed on a level sandy surface is 4 mph instead of the desired 7 mph. This is due to a lack of necessary power at the wheels. The lack of power resulted from dimension considerations that allowed only an eight horsepower engine and also from mechanical inefficiencies of the hydraulic system. However, the vehicle will be able to climb a 30 percent grade, surpass a one meter high object and one meter wide crevice. The wheel assemblies will be able to support the entire weight of the vehicle on two wheels. The wheels will also provide adequate suspension for the vehicle and sufficient traction for lunar-type surfaces.

Nutt, Richard E.↗
NASA NTRS2019

Mechanical Design and Configuration of Penetrations for the Europa Clipper Avionics Vault Structure

The main purpose of the Avionics Vault is to shield radiation sensitive electronics for the Europa Clipper Spacecraft. The vault is a box structure made out of aluminum panels. The panels are roughly 10 mm thick in order to shield the electronics from the orbital total ionizing radiation around Jupiter. The vault requires an electromagnetic interference (EMI) shielding effectiveness (SE) of at least 70 dB in order to mitigate EMI with the spacecraft radar receiver. Overall, the vault accommodates four main types of penetrations: receptacle connectors, pass-through cables, fluid lines, and vent holes. More than 150 cables penetrate the vault panels to connect to electronic boxes inside. Fluid pipes enter and exit the vault to transfer heat to the rest of the spacecraft. Vent holes provide a path for air to escape from the vault during launch. Several novel penetrations designs were created to meet EMI and radiation shielding requirements. Receptacle connectors interface to the vault panels using 1.3 mm thick Ta10W plates. Pass-through cables penetrate the vault using aluminum clamshells after being wrapped with Teflon cushion tape, Kapton tape, and copper tape. Vent hole penetrations consist of a copper mesh for EMI shielding and an aluminum radiation shield bracket to direct air out of the vault during launch. Fluid lines terminate at the vault wall using mechanical fittings that resemble a nut and bolt interface. In addition, most mechanical seams and penetrations utilize EMI gaskets to ensure proper EMI shielding. To reduce risk and confirm that the vault penetration designs were appropriate for EMI shielding, an EMI chamber at the Jet Propulsion Laboratory (JPL) was used to test a mock-up vault panel with multiple variations of all four main types of vault penetrations. This EMI SE test also incorporated different methods for bundling pass-through cables, and a comparison of flange mounted connectors versus jam nut connectors. A low noise preamplifier and a Rohde & Schwarz spectrum analyzer measured E-field levels transmitting through the mock-up vault panel. The results showed a shielding effectiveness of 77 dB for the mock-up vault panel, which exceeds the 70 dB target for Europa Clipper. Both the flange mounted connectors and jam nut connectors exhibited similar EMI SE results at the measured frequencies, and all variations of vault penetrations showed favorable EMI SE levels. Since the flight panels will be much larger and include many more penetrations, there will be testing of the flight vault to confirm its EMI SE is compliant with environmental requirements.

Giersch, Louis↗
NASA NTRS1988

Mechanical design engineering. NASA/university advanced design program: Lunar Bulk Material Transport Vehicle

The design of a Lunar Bulk Material Transport Vehicle (LBMTV) is discussed. Goals set in the project include a payload of 50 cubic feet of lunar soil with a lunar of approximately 800 moon-pounds, a speed of 15 mph, and the ability to handle a grade of 20 percent. Thermal control, an articulated steering mechanism, a dump mechanism, a self-righting mechanism, viable power sources, and a probable control panel are analyzed. The thermal control system involves the use of small strip heaters to heat the housing of electronic equipment in the absence of sufficient solar radiation and multi-layer insulation during periods of intense solar radiation. The entire system uses only 10 W and weighs about 60 pounds, or 10 moon-pounds. The steering mechanism is an articulated steering joint at the center of the vehicle. It utilizes two actuators and yields a turning radius of 10.3 feet. The dump mechanism rotates the bulk material container through an angle of 100 degree using one actuator. The self-righting mechanism consists of two four bar linkages, each of which is powered by the same size actuator as the other linkages. The LBMTV is powered by rechargeable batteries. A running time of at least two hours is attained under a worst case analysis. The weight of the batteries is 100 pounds. A control panel consisting of feedback and control instruments is described. The panel includes all critical information necessary to control the vehicle remotely. The LBMTV is capable of handling many types of cargo. It is able to interface with many types of removable bulk material containers. These containers are made to interface with the three-legged walker, SKITTER. The overall vehicle is about 15 feet in length and has a weight of about 1000 pounds, or 170 lunar pounds.

Daugherty, Paul↗
NASA NTRS1992

Mechanical design of a rotary balance system for NASA. Langley Research Center's vertical spin tunnel

A new, lightweight Rotary Balance System is presently being fabricated and installed as part of a major upgrade to the existing 20-Foot Vertical Spin Tunnel (Figure 1) located at NASA's Langley Research Center. This upgrade, to improve model testing productivity of the only free-spinning vertical wind tunnel in the Western Hemisphere, includes a modern fan/drive and tunnel control system, an updated video recording system, and the new rotary balance system. The rotary balance is a mechanical apparatus which enables the measurement of aerodynamic force and moment data under spinning conditions (100 rpm). This data is used in spin analysis and is vital to the implementation of large amplitude maneuvering simulations required for all new high performance aircraft. The present system is over 40 years old and is worn beyond acceptable tolerances, as well as being susceptible to frequent breakdowns. It has serious limitations in model weight, position adjustment and spin-rate capabilities. The new rotary balance system described in this report will permit greater test efficiency and improved data accuracy. Rotary Balance testing with the model enclosed in a tare bag can also be performed to obtain resulting model forces from the spinning operation. The rotary balance system will be stored against the tunnel sidewall during free-flight model testing.

J. W. Allred↗
DOE OSTI2025

Mechanical Actuation Mechanism Design for Cryogenic Microwave Cavities in Axion Dark Matter Searches

Resonant Microwave Cavities are placed in a strong magnetic field at cryogenic temperatures (~.1K) in order to try to convert axions into detectable photons. A major challenge is to tune the cavity to allow no interference with the desired environment used for the experiment, which enables the system to operate near absolute zero and maintain superconductivity, thereby reducing noise as much as possible. My project focuses on developing and testing a low-thermal-conductivity gear reduction system with a 1000:1 ratio for cavity tuning. This gear system could significantly improve the thermal and electromagnetic interference of the tuning mechanism, helping ADMX scan for axions more accurately.

Ortega, Giancarlo [Marquette U.; Fermilab]↗
DOE OSTI2025

Mechanical Design and Vacuum Pressure Impregnation of Combined Function Electromagnets

There are 96 combined function magnets, 48 focusing and 48 defocusing, that have been in operation in the Fermi NAL Booster without a critical failure requiring replacement for over 50 years. The Proton Improvement Plan-II (PIP-II) poses risk to these magnets and Fermilab does not currently have the tooling or knowledge of process to fabricate spares. Two new variants of these magnets are needed to reduce beam loss by having a larger aperture (Booster Gradient Defocusing Wide, or BGDS) and create space for a new injection system (Booster Gradient Defocusing Short, or BGDS). Risk of failure and need for spares will increase substantially with the implementation of PIP-II in consequence of a frequency increase from 15 Hz to 20 Hz and a 30 percent increase in peak to ground voltage. The BGDS and BGDW magnets will be fabricated in the same way as the current booster magnets so that future combined function magnet spares can be produced using learned processes.

Szabo, B. [Fermilab]↗