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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS2005

In-Space Cryogenic Propellant Depot Stepping Stone

An In-Space Cryogenic Propellant Depot (ISCPD) is an important stepping stone to provide the capability to preposition, store, manufacture, and later use the propellants for Earth-Neighborhood campaigns and beyond. An in-space propellant depot will provide affordable propellants and other similar consumables to support the development of sustainable and affordable exploration strategies as well as commercial space activities. An in-space propellant depot not only requires technology development in key areas such as zero boil-off storage and fluid transfer, but in other areas such as lightweight structures, highly reliable connectors, and autonomous operations. These technologies can be applicable to a broad range of propellant depot concepts or specific to a certain design. In addition, these technologies are required for spacecraft and orbit transfer vehicle propulsion and power systems, and space life support. Generally, applications of this technology require long-term storage, on-orbit fluid transfer and supply, cryogenic propellant production from water, unique instrumentation and autonomous operations. This paper discusses the reasons why such advances are important to future affordable and sustainable operations in space. This paper also discusses briefly R&D objectives comprising a promising approach to the systems planning and evolution into a meaningful stepping stone design, development, and implementation of an In-Space Cryogenic Propellant Depot. The success of a well-planned and orchestrated approach holds great promise for achieving innovation and revolutionary technology development for supporting future exploration and development of space.

Howell, Joe T.↗
NASA NTRSDate not supplied

Magnetohydrodynamics (MHD) Aerocapture System for Enabling Faster-Larger Planetary Science & Human Exploration Missions

Since our completing the NIAC Phase I NIAC on this Advanced Aerocapture System, NASA Langley Research Center has funded or supported a number of studies and code enhancements through its Center Innovation Fund (CIF) and NASA’s NSTGRO and Internship Programs to mature the analysis capabilities and quantify the merits of the MHD Aerocapture System technology. These efforts have resulted in a plug and play analysis capability for assessing MHD aerocapture system performance for arrival at many planetary bodies of interest. Our efforts have especially focused on the potential mass savings for improving the capacity for science observations at Neptune and Triton. A re-cent Forbes article published “‘Orbital mechanics is probably going to decide for us whether we go to Uranus or Neptune because we need to flyby Jupiter,’ said Kunio Sayanagi at Hampton University, Virginia, who also worked on the Neptune Odyssey proposal…. Exactly when a mission can be sent to Uranus, or Neptune, depends on the relative position of Jupiter, which can help give a spacecraft a gravitational slingshot. That drastically shortens the cruise phase.” [1] Since shortening the cruise phase is important for these science missions, any mass savings enabled by the MHD Aerocapture System could be reallocated to increasing Thermal Protection System mass to allow faster arrival speeds and/or for onboarding additional payloads such as science instruments, batteries, or propellant for conducting more science for longer durations in the desired orbits. The analysis steps and codes for conducting trades and sizing vehicles for aerocapture are as follows: Step 1 is to conduct aeroheating analysis using LAURA of the selected entry vehicle shape to identify locations on the forebody where ionization and flow velocity are sufficient for producing Lo-rentz forces. LAURA is a multiblock structured grid finite-volume CFD solver developed at the NASA Langley Research Center. [2] LAURA has been used for aerothermal analysis support of the entry, de-scent and landing (EDL) phase of interplanetary missions over the last three decades [3-7]. Step 2 is to port the LAURA results into CFDWARP to calcu-late electrical and thermal conductivities of ionized flow for sizing MHD patch system and calculating Lorentz forces needed for controls analysis. CFDWARP is a CFD code that uses advanced nu-merical methods that enable the simulation of the full coupling between the aerodynamics, the magne-tohydrodynamics, and the non-neutral plasma sheaths. CFDWARP has the unique capability to simulate efficiently the non-neutral sheaths (near the electrodes) in coupled form with the quasi-neutral bulk MHD flow [8-11]. Step 3 is to link re-sults from LAURA and CFDWARP into POST2 for calculating entry trajectories and comparing MHD control results with other aerodynamic control strategies. The Program to Optimize Simulated Tra-jectories II (POST2) is a generalized point mass, discrete parameter targeting and optimization pro-gram. POST2 provides the capability to target and optimize point mass trajectories for multiple pow-ered or un-powered vehicles near an arbitrary rotat-ing, oblate planet [12]. Step 4: TPS sizing was per-formed using the Fully Implicit Ablation and Ther-mal-response code (FIAT) tool which computes the transient one-dimensional thermal response and surface thermochemistry of a multilayer stackup of thermal protection, bonding, and structural materi-als subject to aeroheating on one surface [13]. The sizing and margining methodology used was based on the approach documented by Mahzari and Milos [14] for the dual-layered heatshield for extreme entry environment technology (DL-HEEET) TPS concept. TPS analysis utilizes trajectory information from POST2. Using this step-wise plug and play MHD Aerocapture performance assessment process, our analysis targets a Neptune aerocapture trajectory that will place the spacecraft in an observation orbit for Triton. [15]. Magnetohydrodynamic (MHD) control of a 4.5-meter diameter MSL-style capsule resulted in TPS mass savings of nearly 2000 kg when using an MHD system mass of under 200 kg. The flight path for a vehicle using the MHD control strategy has a much lower heat rate and heat load compared to the conventional aerodynamic aerocapture strategies known as bank angle con-trolled (BAC) and direct force controlled (DFC). Both BAC and DFC have heat rates significantly greater than 1500 W/cm2 typically used as an upper limit for PICA. Thus, DL-HEEET TPS concept was required for the BAC and DFC control strategies. However, considering the more benign environ-ments for the MHD case, additional TPS concepts with improved mass efficiency were also assessed. PICA was considered for the MHD controlled strat-egy since the maximum heat rate was well within the limits (<1500 W/cm2) of PICA. TPS sizing re-sulted in a significant mass reduction. The PICA layer for this sizing case was about 7.8 cm. As a point of reference, the Mars 2020 mission, which used this same PICA concept, had a PICA thickness of 3.18 cm [16]. The trajectories used for the TPS sizing originat-ed from the POST2 simulations. The current, I, to an electromagnet configuration can be manipulated to allow for active control of the vehicle. Manipula-tion of the current, I, changes the magnetic field, B, which affects the Lorentz force and therefore the MHD drag force on the vehicle. Our analysis in-cluded both open-loop and close-loop control. Closed-loop control will enable improved overall performance when taking into account mission level uncertainties, such as interplanetary delivery errors and atmospheric modeling uncertainties. The open-loop and closed-loop MHD control cases do not dip as deep into the atmosphere as the aerodynamic cases. Three types of aerodynamic-only approach-es are investigated: bank angle modulation (BAM), director force control (DFC), and Drag Modulated. BAM and DFC make use of vehicle aerodynamic angles to steer the vehicle. Thus, changing the aer-odynamic forces acting on the vehicle for control, aerodynamic drag modulated case requires a vary-ing drag area to modulate the drag force. The MHD drag modulated case modulates MHD generated drag force that adds to the aerodynamic drag. This higher atmospheric activation of drag forces by the MHD patch results in significantly less heat flux on the vehicle. The MHD technology will enable shorter cruise times and deceleration of larger payloads for increasing the capacity for science at the Ice Giants or for returning astronauts to Earth from cislunar space or from Mars. The purpose of this presentation is to provide more details about this work and to highlight plans for further research and development including a flight demonstration.

R. W. Moses↗
NASA NTRS1971

The combustion of pure and composite propellants - The expansion and application of laminar flame theory to heterogeneous solid propellants.

Development of a combustion theory based on the laminarized solutions to the energy and flow conservation equations, which is more realistic in recognizing the nature of the heating-rate problem and in obtaining a practical solution to estimating its magnitude. A new experimental approach is used for studying the combustion behavior of pure monopropellants and composite propellants which uses a laser beam to supply additional heat feedback to a burning surface. New experimental data are presented for the laser-induced combustion rate and ignition delay of pure ammonium perchlorate. The pure monopropellant theory is generalized to include such nonadiabatic effects, and the new experimental data are in good agreement with the nonadiabatic theory.-

Hertzberg, M.↗
NASA NTRS1974

RF gauging efforts with liquid hydrogen and liquid oxygen as applicable to subcritical space vehicle systems

The RF gauging concept is based on the interaction between a fluid dielectric medium in an enclosed metallic cavity and electromagnetic fields set up within that cavity. In RF gauging systems, the fundamental measurement relies on the interpretation of changes in the resonant RF frequencies of an enclosed tank as the mass of the propellant contained in the tank is changed. In addition to a discussion of the basic principles of operation of these systems, the study presents a description of the current breadboard implementation with typical test arrangements, along with supporting test data. The experimental testing of the RF gauging technique for liquid cryogen mass gauging indicates that this technique is a feasible approach to liquid oxygen and liquid hydrogen gauging under all attitude or reduced gravity environments.

Thompson, H. E.↗
NASA NTRS1988

Manned Mars mission accommodation: Sprint mission

The results of a study conducted at the NASA-LaRC to assess the impacts on the Phase 2 Space Station of Accommodating a Manned Mission to Mars are documented. In addition, several candidate transportation node configurations are presented to accommodate the assembly and verification of the Mars Mission vehicles. This study includes an identification of a life science research program that would need to be completed, on-orbit, prior to mission departure and an assessment of the necessary orbital technology development and demonstration program needed to accomplish the mission. Also included is an analysis of the configuration mass properties and a preliminary analysis of the Space Station control system sizing that would be required to control the station. Results of the study indicate the Phase 2 Space Station can support a manned mission to Mars with the addition of a supporting infrastructure that includes a propellant depot, assembly hangar, and a heavy lift launch vehicle to support the large launch requirements.

Cirillo, William M.↗
NASA NTRS1992

Oxygen production on the Lunar materials processing frontier

During the pre-conceptual design phase of an initial lunar oxygen processing facility, it is essential to identify and compare the available processes and evaluate them in order to ensure the success of such an endeavor. The focus of this paper is to provide an overview of materials processing to produce lunar oxygen as one part of a given scenario of a developing lunar occupation. More than twenty-five techniques to produce oxygen from lunar materials have been identified. While it is important to continue research on any feasible method, not all methods can be implemented at the initial lunar facility. Hence, it is necessary during the pre-conceptual design phase to evaluate all methods and determine the leading processes for initial focus. Researchers have developed techniques for evaluating the numerous proposed methods in order to suggest which processes would be best to go to the Moon first. As one section in this paper, the recent evaluation procedures that have been presented in the literature are compared and contrasted. In general, the production methods for lunar oxygen fall into four categories: thermochemical, reactive solvent, pyrolytic, and electrochemical. Examples from two of the four categories are described, operating characteristics are contrasted, and terrestrial analogs are presented when possible. In addition to producing oxygen for use as a propellant and for life support, valuable co-products can be derived from some of the processes. This information is also highlighted in the description of a given process.

Altenberg, Barbara H.↗
NASA NTRS2000

POWOW: A Modular, High Power Spacecraft Concept

A robust space infrastructure encompasses a broad range of mission needs along with an imperative to reduce costs of satellites meeting those needs. A critical commodity for science, commercial and civil satellites is power at an affordable cost. The POWOW (POwer WithOut Wires) spacecraft concept was created to provide, at one end of the scale, multi-megawatts of power yet also be composed of modules that can meet spacecraft needs in the kilowatt range. With support from the NASA-sponsored Space Solar Power Exploratory Research and Technology Program, the POWOW spacecraft concept was designed to meet Mars mission needs - while at the same time having elements applicable to a range of other missions. At Mars, the vehicle would reside in an aerosynchronous orbit and beam power to a variety of locations on the surface. It is the purpose of this paper to present the latest concept design results. The Space Power Institute along with four companies: Able Engineering, Inc., Entech, Inc., Primex Aerospace Co., and TECSTAR have produced a modular, power-rich electrically propelled spacecraft design that meets these requirements. In addition, it also meets a range of civil and commercial needs. The spacecraft design is based on multijunction Ill-V solar cells, the new Stretched Lens Aurora (SLA) module, a lightweight array design based on a multiplicity of 8 kW end-of-life subarrays and electric thrusters. The solar cells have excellent radiation resistance and efficiencies above 30%. The SLA has a concentration ratio up to 15x while maintaining an operating temperature of 80 C. The design of the 8 kW array building block will be presented and its applicability to commercial and government missions will be discussed. Electric propulsion options include Hall, MPD and ion thrusters of various power levels and trade studies have been conducted to define the most advantageous options. The present baseline spacecraft design providing 900 kW using technologies expected to be available in 2003 will be described. Areal power densities of nearly 400 W/meters squared at 80 C operating temperatures and wing level specific powers of over 400 W/kg are projected. Details of trip times and payloads to Mars will be presented as well as trade studies of various electric propulsion options. Trip times compare favorably with chemical propulsion options. Because the design is modular, learning curve methodology can be applied to determine expected cost reductions. These results will also be included. This paper has not been presented at a previous meeting.

Brandhorst, Henry W., Jr.↗
NASA NTRS2000

Friction Plug Weld Repair for the Space Shuttle External Tank

Lockheed Martin Space Systems, Michoud Operations in New Orleans, LA is the manufacturer of the External Fuel Tanks (ET) for the Space Transportation System (STS). The ET contains and delivers the propellants used by the Orbiters three main engines. Additionally, it also serves as the structural backbone for the Orbiter and the two Solid Rocket Boosters (SRB), which combined, constitute the STS. In 1994, NASA established that in order to launch the International Space Station, the performance of the STS must be improved. One option was to reduce the weight of the ET, which would enable sufficient increase in performance. With the development of the Weldalite(R) series of Al-Cu-Li alloys in the late 1980's, Lockheed Martin was postured to replace the current A12219 fuel tanks with the high strength, light weight A12195 alloy. With the use of A12195 and some component redesign, the weight of the Super Lightweight (SLWT) ET was reduced by approximately 7,000 pounds, which added as much capability to the Space Shuttle. Since June 1998, seven STS missions have been successful with the use of the SLWT ET's.

Hartley, Paula J.↗
NASA NTRS2000

A Survey and Recent Development of Lunar Gravity Assist

Earth's moon is the largest in the solar system relative to its parent body, the Earth, and can have significant effect on the path of a spacecraft flying close by. This effect, when planned to benefit a specific mission, is called lunar gravity assist (LGA) , and assumes that one aims the spacecraft towards the Moon in such a way that the Moon's gravitational pull will alter the spacecraft's course in a favorable manner. The first application of LGA was in the Apollo program, where the command and lunar modules (and astronauts) were propelled to the Moon such that, if no additional course changes were made, they would swing around the backside of the Moon at a certain altitude and be flung back to Earth to enter the atmosphere at a specified location in the Pacific ocean. This LGA was an essential element saving the lives of the astronauts on Apollo 13. This paper will illustrate the basic mechanics of gravity assist, and list the many applications where it has been used effectively over the past 30 some years. These include missions to the sun and Earth libration points, redirecting a spacecraft from one of these point to a comet encounter, and enhancing payloads by providing an energy boost by the Moon. More recently, studies and actual missions have shown the benefits of LGA in: (1) assisting lunar capture, (2) repositioning geosynchronous communications satellites, (3) boosting spacecraft to Earth escape and departure to planets and other solar system bodies, and (4) allowing small spacecraft to be launched as secondary payloads and released into almost a random orbit from which each may depart and maneuver in space with gravity assists from the Earth and Moon to perform a specific planetary or other mission. This latter application is a recent development by the author and is being applied in 2002 and later years, with piggyback flights on the Ariane 5 which launches comsats to GEO.

Penzo, Paul A.↗
NASA NTRS2003

Enhancement and Extension of Porosity Model in the FDNS-500 Code to Provide Enhanced Simulations of Rocket Engine Components

In the past, the design of rocket engines has primarily relied on the cold flow/hot fire test, and the empirical correlations developed based on the database from previous designs. However, it is very costly to fabricate and test various hardware designs during the design cycle, whereas the empirical model becomes unreliable in designing the advanced rocket engine where its operating conditions exceed the range of the database. The main goal of the 2nd Generation Reusable Launching Vehicle (GEN-II RLV) is to reduce the cost per payload and to extend the life of the hardware, which poses a great challenge to the rocket engine design. Hence, understanding the flow characteristics in each engine components is thus critical to the engine design. In the last few decades, the methodology of computational fluid dynamics (CFD) has been advanced to be a mature tool of analyzing various engine components. Therefore, it is important for the CFD design tool to be able to properly simulate the hot flow environment near the liquid injector, and thus to accurately predict the heat load to the injector faceplate. However, to date it is still not feasible to conduct CFD simulations of the detailed flowfield with very complicated geometries such as fluid flow and heat transfer in an injector assembly and through a porous plate, which requires gigantic computer memories and power to resolve the detailed geometry. The rigimesh (a sintered metal material), utilized to reduce the heat load to the faceplate, is one of the design concepts for the injector faceplate of the GEN-II RLV. In addition, the injector assembly is designed to distribute propellants into the combustion chamber of the liquid rocket engine. A porosity mode thus becomes a necessity for the CFD code in order to efficiently simulate the flow and heat transfer in these porous media, and maintain good accuracy in describing the flow fields. Currently, the FDNS (Finite Difference Navier-Stakes) code is one of the CFD codes which are most widely used by research engineers at NASA Marshall Space Flight Center (MSFC) to simulate various flow problems related to rocket engines. The objective of this research work during the 10-week summer faculty fellowship program was to 1) debug the framework of the porosity model in the current FDNS code, and 2) validate the porosity model by simulating flows through various porous media such as tube banks and porous plate.

Cheng, Gary↗
NASA NTRS2004

Restoring Redundancy to the Wilkinson Microwave Anisotrophy Probe Propulsion System

The Wilkinson Microwave Anisotropy Probe is a follow-on to the Differential Microwave Radiometer instrument on the Cosmic Background Explorer. Attitude control system engineers discovered sixteen months before launch that configuration changes after the critical design review had resulted in a significant migration of the spacecraft's center of mass. As a result, the spacecraft no longer had a viable backup control mode in the event of a failure of the negative pitch-axis thruster. A tiger team was formed and identified potential solutions to this problem, such as adding thruster-plume shields to redirect thruster torque, adding or removing mass from the spacecraft, adding an additional thruster, moving thrusters, bending thruster nozzles or propellant tubing, or accepting the loss of redundancy. The project considered the impacts on mass, cost, fuel budget, and schedule for each solution, and decided to bend the propellant tubing of the two roll-control thrusters to allow the pair to be used for backup control in the negative pitch axis. This paper discusses the problem and the potential solutions, and documents the hardware and software changes and verification performed. Flight data are presented to show the on-orbit performance of the propulsion system and lessons learned are described.

O'Donnell, James R., Jr.↗
NASA NTRS2005

Cassini - Huygens maneuver experience : cruise and arrival at Saturn

The Cassini-Huygens mission to Saturn and Titan was launched in 1997. It is an international effort to study the Saturnian system. Cassini-Huygens' interplanetary cruise delivered the spacecraft to Saturn in 2004. It also made use of many propulsive maneuvers, both statistical and deterministic. Maneuver-related analysis and performance for latter half of cruise is reported. The system has performed more accurately than the pre-launch expectations and requirements. Additionally, some maneuvers have already been skipped, saving propellant and flight team effort. Analysis of historical execution error data is presented.

Huygens↗
NASA NTRS2010

Life Support with Failures and Variable Supply

The life support system for long duration missions will recycle oxygen and water to reduce the material resupply mass from Earth. The impact of life support failures was investigated by dynamic simulation of a lunar outpost habitat life support model. The model was modified to simulate resupply delays, power failures, recycling system failures, and storage failures. Many failures impact the lunar outpost water supply directly or indirectly, depending on the water balance and water storage. Failure effects on the water supply are reduced if Extra Vehicular Activity (EVA) water use is low and the water supply is ample. Additional oxygen can be supplied by scavenging unused propellant or by production from regolith, but the amounts obtained can vary significantly. The requirements for oxygen and water can also vary significantly, especially for EVA. Providing storage buffers can improve efficiency and reliability, and minimize the chance of supply failing to meet demand. Life support failures and supply variations can be survivable if effective solutions are provided by the system design

Jones, Harry↗
NASA NTRS2015

The Iodine Satellite (iSat) Project Development Towards Critical Design Review

Despite the prevalence of small satellites in recent years, the systems flown to date have very limited propulsion capability. SmallSats are typically secondary payloads and have significant constraints for volume, mass, and power in addition to limitations on the use of hazardous propellants or stored energy. These constraints limit the options for SmallSat maneuverability. NASA's Space Technology Mission Directorate approved the iodine Satellite flight project for a rapid demonstration of iodine Hall thruster technology in a 12U (cubesat units) configuration under the Small Spacecraft Technology Program. The mission is a partnership between NASA MSFC, NASA GRC, and Busek Co, Inc., with the Air Force supporting the propulsion technology maturation. The team is working towards the critical design review in the final design and fabrication phase of the project. The current design shows positive technical performance margins in all areas. The iSat project is planned for launch readiness in the spring of 2017.

Dankanich, John W.↗
NASA NTRS2016

The Iodine Satellite (iSat) Project Development Towards Critical Design Review (CDR)

Despite the prevalence of Small Satellites in recent years, the systems flown to date have very limited propulsion capability. SmallSats are typically secondary payloads and have significant constraints for volume, mass, and power in addition to limitations on the use of hazardous propellants or stored energy (i.e. high pressure vessels). These constraints limit the options for SmallSat maneuverability. NASA's Space Technology Mission Directorate approved the iodine Satellite flight project for a rapid demonstration of iodine Hall thruster technology in a 12U configuration under the Small Spacecraft Technology Program. The project formally began in FY15 as a partnership between NASA MSFC, NASA GRC, and Busek Co, Inc., with the Air Force supporting the propulsion technology maturation. The team is in final preparation of the Critical Design Review prior to initiating the fabrication and integration phase of the project. The iSat project is on schedule for a launch opportunity in November 2017.

Electric Propulsion↗
NASA NTRS2018

Electric Motor Noise for Small Quadcopters: Part I - Acoustic Measurements

There is increased interest in using electric motors to drive propulsors across a range of small air vehicle classes. Applications include both vertical lift and conventional takeoff and landing systems for Small Unmanned Aircraft Systems. Mission profiles call for integrating these systems into urban airspaces exposing populated areas to new noise sources. In addition to the propulsor noise from rotors and propellers, electric motors are expected to contribute to the overall sound levels and possibly human annoyance. This work presents acoustic measurements of electric motors used for small quadcopters to characterize the sound and identify sources with and without a propeller. Free field microphone measurements were used to determine directivity and a phased microphone array was used to identify sound sources. A companion paper (Part II – Source Characteristics and Prediction) compares the far field results with current probe measurements of the signal driving the motor, the structural response of the motor case, and describes prediction methods of electric motor noise.

Aircraft Noise↗
NASA NTRS2018

Electric Motor Noise from Small Quadcopters: Part 1 - Acoustic Measurements

There is increased interest in using electric motors to drive propulsors across a range of small air vehicle classes. Applications include both vertical lift and conventional takeoff and landing systems for Small Unmanned Aircraft Systems. Mission profiles call for integrating these systems into urban airspaces exposing populated areas to new noise sources. In addition to the propulsor noise from rotors and propellers, electric motors are expected to contribute to the overall sound levels and possibly human annoyance. This work presents acoustic measurements of electric motors used for small quadcopters to characterize the sound and identify sources with and without a propeller. Free field microphone measurements were used to determine directivity and a phased microphone array was used to identify sound sources. A companion paper (Part II - Source Characteristics and Prediction) compares the far field results with current probe measurements of the signal driving the motor, the structural response of the motor case, and describes prediction methods of electric motor noise.

Electric Propulsion↗