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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 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Heating Analysis of the Lockheed Lifting Body

Computational fluid dynamic (CFD) analysis is performed on the Lockheed Lifting Body Single-Stage-to-Orbit vehicle to determine the heat transfer to the vehicle during its descent trajectory. Seven species, chemical nonequilibriurn computations using the GASP code will be completed at several trajectory points to assess the thermal protection requirements of the vehicle. Sophisticated surface boundary conditions including in-depth conduction, catalycity, and a variable temperature wall have been incorporated into the flow solver.

Palmer, Grant↗
NASA NTRS2008

Constraint Force Equation Methodology for Modeling Multi-Body Stage Separation Dynamics

This paper discusses a generalized approach to the multi-body separation problems in a launch vehicle staging environment based on constraint force methodology and its implementation into the Program to Optimize Simulated Trajectories II (POST2), a widely used trajectory design and optimization tool. This development facilitates the inclusion of stage separation analysis into POST2 for seamless end-to-end simulations of launch vehicle trajectories, thus simplifying the overall implementation and providing a range of modeling and optimization capabilities that are standard features in POST2. Analysis and results are presented for two test cases that validate the constraint force equation methodology in a stand-alone mode and its implementation in POST2.

Toniolo, Matthew D.↗
NASA NTRS1978

Performance evaluation of integrated guidance and navigation systems

The method of linear covariance analysis is applied to the study of an aerospace vehicle integrated guidance-and-navigation system. This requires that the rms trajectory deviations of the vehicle be computed along with the rms errors of the navigation system. The technique is used to examine the performance of the shuttle orbiter vehicle's guidance-and-navigation system during the terminal phase of the landing. Essentially the same rms trajectory deviations were obtained from a single covariance run as from a 30-sample Monte-Carlo run set. Of particular interest in the study was the result that rms trajectory deviations for the integrated-system were generally significantly larger than the rms navigation errors.

Tao, Y. C.↗
NASA NTRS2000

Thermal Analysis Methods for an Earth Entry Vehicle

Thermal analysis of a vehicle designed to return samples from another planet, such as the Earth Entry vehicle for the Mars Sample Return mission, presents several unique challenges. The Earth Entry Vehicle (EEV) must contain Martian material samples after they have been collected and protect them from the high heating rates of entry into the Earth's atmosphere. This requirement necessitates inclusion of detailed thermal analysis early in the design of the vehicle. This paper will describe the challenges and solutions for a preliminary thermal analysis of an Earth Entry Vehicle. The aeroheatina on the vehicle during entry would be the main driver for the thermal behavior. and is a complex function of time, spatial position on the vehicle, vehicle temperature, and trajectory parameters. Thus. the thermal analysis must be closely tied to the aeroheating analysis in order to make accurate predictions. Also, the thermal analysis must account for the material response of the ablative thermal protection system TPS. For the exo-atmospheric portion of the mission, the thermal analysis must include the orbital radiation fluxes on the surfaces. The thermal behavior must also be used to predict the structural response of the vehicle (the thermal stress and strains) and whether they remain within the capability of the materials. Thus, the thermal analysis requires ties to the three-dimensional geometry, the aeroheating analysis, the material response analysis, the orbital analysis. and the structural analysis. The goal of this paper is to describe to what degree that has been achieved.

Amundsen, Ruth M.↗
NASA NTRS2000

Thermal Analysis Methods For Earth Entry Vehicle

Thermal analysis of a vehicle designed to return samples from another planet, such as the Earth Entry vehicle for the Mars Sample Return mission, presents several unique challenges. The Earth Entry Vehicle (EEV) must contain Martian material samples after they have been collected and protect them from the high heating rates of entry into the Earth's atmosphere. This requirement necessitates inclusion of detailed thermal analysis early in the design of the vehicle. This paper will describe the challenges and solutions for a preliminary thermal analysis of an Earth Entry Vehicle. The aeroheating on the vehicle during entry would be the main driver for the thermal behavior, and is a complex function of time, spatial position on the vehicle, vehicle temperature, and trajectory parameters. Thus, the thermal analysis must be closely tied to the aeroheating analysis in order to make accurate predictions. Also, the thermal analysis must account for the material response of the ablative thermal protection system (TPS). For the exo-atmospheric portion of the mission, the thermal analysis must include the orbital radiation fluxes on the surfaces. The thermal behavior must also be used to predict the structural response of the vehicle (the thermal stress and strains) and whether they remain within the capability of the materials. Thus, the thermal analysis requires ties to the three-dimensional geometry, the aeroheating analysis, the material response analysis, the orbital analysis, and the structural analysis. The goal of this paper is to describe to what degree that has been achieved.

Amundsen, Ruth M.↗
NASA NTRS2015

An Examination of "The Martian" Trajectory

This analysis was performed to support a request to examine the trajectory of the Hermes vehicle in the novel "The Martian" by Andy Weir. Weir developed his own tool to perform the analysis necessary to provide proper trajectory information for the novel. The Hermes vehicle is the interplanetary spacecraft that shuttles the crew to and from Mars. It is notionally a Nuclear powered vehicle utilizing VASIMR engines for propulsion. The intent of this analysis was the determine whether the trajectory as it was outlined in the novel is consistent with the rules of orbital mechanics.

Mars Trajectory Analysis↗
NASA NTRS2002

Improvements in Thermal Protection Sizing Capabilities for TCAT: Conceptual Design for Advanced Space Transportation Systems

The Thermal Calculation Analysis Tool (TCAT), originally developed for the Space Systems Design Lab at the Georgia Institute of Technology, is a conceptual design tool capable of integrating aeroheating analysis into conceptual reusable launch vehicle design. It provides Thermal Protection System (TPS) unit thicknesses and acreage percentages based on the geometry of the vehicle and a reference trajectory to be used in calculation of the total cost and weight of the vehicle design. TCAT has proven to be reasonably accurate at calculating the TPS unit weights for in-flight trajectories; however, it does not have the capability of sizing TPS materials above cryogenic fuel tanks for ground hold operations. During ground hold operations, the vehicle is held for a brief period (generally about two hours) during which heat transfer from the TPS materials to the cryogenic fuel occurs. If too much heat is extracted from the TPS material, the surface temperature may fall below the freezing point of water, thereby freezing any condensation that may be present at the surface of the TPS. Condensation or ice on the surface of the vehicle is potentially hazardous to the mission and can also damage the TPS. It is questionable whether or not the TPS thicknesses provided by the aeroheating analysis would be sufficiently thick to insulate the surface of the TPS from the heat transfer to the fuel. Therefore, a design tool has been developed that is capable of sizing TPS materials at these cryogenic fuel tank locations to augment TCAT's TPS sizing capabilities.

Olds, John R.↗
NASA NTRSDate not supplied

Design and Technology Maturation of the Stratospheric Projectile Experiment of Entry Dynamics

The supersonic and transonic dynamic stability of blunt-body reentry vehicles currently poses large risks in all of NASA’s ongoing entry missions (MSR SRL, MSR EES, and Dragonfly). These projects have allocated millions of dollars to testing and modeling efforts to buy down risk by using the current state-of-the-art (SoA) facilities at NASA’s disposal. While these facilities have heritage in supplying dynamics data to reentry missions, their availability is severely limited – particularly with the high number of concur-rent projects requesting simultaneous testing– and are costly when considering the science density per dollar. None of the current SoA facility methodologies allow the test model to have the dynamics fully develop through a flight relevant free-stream profile and as such require extrapolations with resultant high uncertainties in order to relate the test dynamics to flight expectations. SPEED is a NASA Ames Center Innovation Fund (CIF) project that is developing a highly tailorable and cost-effective test methodology to better assess the dynamic stability of blunt-body reentry vehicles via a stratospheric balloon flight. This is accomplished by dropping a suite of instrumented capsules from a stratospheric balloon to gain a statistically relevant dataset of scaled reentry vehicles in mission relevant free-flight conditions. This presentation will walk through how the test methodology is being implemented specifically for the Mars Sample Return (MSR) Earth Entry System (EES) geometry in an awarded Flight Opportunities Program (FOP) test flight in early CY24. SPEED Application to MSR: SPEED consists of three main mechanical systems: the Drop Platform, the Projectile, and the test Capsule. SPEED is being developed as a set of guidelines and recommendations for how to test with the proposed Concept of Operations (Conops) since the specific design parameters will vary depending on the specific project’s reference trajectory and entry vehicle design. As such, this presentation will walk through the development time-line as shown in Fig. 2. This is meant to serve as a blueprint for further missions as desired. Mechanical and Avionics Design. The SPEED test platform designed for the MSR-EES capsule geometry with nominal entry parameters has the ability to carry 10 Capsules to altitude instrumented with: 1. 3-Axis Accelerometer 2. IMU 3. Gyroscope 4. Magnetometer 5. Pressure Transducer cruciform 6. Uplook and Horizon Cameras To package the avionics/instrumentation suite, the capsule is approximately 1’ in diameter with the Outer Mold Line (OML) centroid-scaled from the full EES design. The internal volume is gutted and custom-shaped to fit the desired instrumentation suite, as well as to allow for the positioning of ballast mass such that the Center of Gravity is analogous to the flight vehicle. All structural components in the Capsule and Projectile are 3D printed, which significantly reduces the cost of each flight unit to around $1500 including all instrumentation, avionics, and structural components. Flight Conops. The test Capsule is accelerated to the desired altitude and Mach number while stowed in the Projectile, a missile-like vehicle consisting of steel ballast in the nose, a low-drag OML, and an Ejection Mechanism to reliably release the Capsule into the free-flow supersonic conditions. For the MSR-EES design, the capsule employs ~3kg of ballast mass at the nose to accelerate the 1.25kg test Capsule to ~Mach 1.7 at 23km altitude. This requires an initial release altitude of 40km, the quoted limit of a 80kg payload by the FOP-contracted balloon provider. Once the Ejection Mechanism avionics detect the proper conditions, the spring-loaded Ejection Mechanism will release and – guided by the sabot – expose the test Capsule to the desired test conditions for ~5 seconds of free-flight in the supersonic/transonic regimes. Dynamics in the subsonic regime will also be captured with the instrumentation suite with post-flight recovery operations aimed at recovering the high-G-load capable SD cards after the planned hard impact landings. Testing and Development: In the few months the SPEED project has worked the development of MSR-EES flight test, the team has performed lab and drone based testing which this presentation will overview. After the first design phase, the team fabricated Engineering Demonstration Units (EDUs) of all subsystems to perform validation testing shown in Fig. 5. After validation was completed on the subsystem level, a drone-drop test was performed at the recreational flight ceiling of 400ft altitude to assess the SPEED systems in a flight environment. Parameters such as in-flight stability, hard impact landing performance, and avionics performance were quantified and qualified. The FY23 CIF will culminate in a helicopter drop test aboard an Air National Guard Blackhawk. This will prepare the team for the CY24 FOP stratospheric balloon flight that should provide the final verification to begin offering the test platform for mission support. Focus of Presentation: This presentation will outline the technology maturation path of the SPEED implementation to the MSR-EES capsule baseline as well as the details regarding the mechanical system, avionics and instrumentation, and flight operations. Note that a complementary presentation is being submitted for a methodology overview of the SPEED test platform, introducing the testing technique and benefits as well as the full application space of the technology.

pitch damping coefficient↗
NASA NTRS2003

High Altitude Launch for a Practical SSTO

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

Landis, Geoffrey A.↗
NASA NTRS2003

High Altitude Launch for a Practical SSTO

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. high-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated in to increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31 % increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

Landis, Geoffrey A.↗
NASA NTRS2003

High Altitude Launch for a Practical SSTO

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2 percent of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20 percent increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2 percent of gross lift-off weight, this corresponds to 31 percent increase in payload (for 5-kilometer launch altitude) to 122 percent additional payload (for 25-kilometer launch altitude).

Landis, Geoffrey A.↗
NASA NTRS2018

Integrated Optimization of Mars Hybrid Solar-Electric/Chemical Propulsion Trajectories

NASA’s Human Exploration and Operation Mission Directorate is developing a reusable hybrid transportation architecture in which both chemical and solar-electric propulsion systems are used to deliver crew and cargo to the Martian sphere of influence. By combining chemical and solar-electric propulsions into a single spacecraft and applying each where it is the most effective, the hybrid architecture enables a series of Mars trajectories that are more fuel efficient than an all chemical propulsion architecture without significant increase to trip time. Solving the complex problem of low-thrust trajectory optimization coupled with the vehicle sizing requires development of an integrated trajectory analysis frame- work. Previous studies have utilized a more segmented optimization framework due to the limitation of the tools available. A new integrated optimization framework was recently developed to address the deficiencies of the previous methods that enables higher fidelity analysis to be performed and increases the efficiency of large design space explorations.

Chai, Patrick R.↗
NASA NTRS2012

Mission Sizing and Trade Studies for Low Ballistic Coefficient Entry Systems to Venus

The U.S and the U.S.S.R. have sent seventeen successful atmospheric entry missions to Venus. Past missions to Venus have utilized rigid aeroshell systems for entry. This rigid aeroshell paradigm sets performance limitations since the size of the entry vehicle is constrained by the fairing diameter of the launch vehicle. This has limited ballistic coefficients (beta) to well above 100 kg/m2 for the entry vehicles. In order to maximize the science payload and minimize the Thermal Protection System (TPS) mass, these missions have entered at very steep entry flight path angles (gamma). Due to Venus thick atmosphere and the steep-gamma, high- conditions, these entry vehicles have been exposed to very high heat flux, very high pressures and extreme decelerations (upwards of 100 g's). Deployable aeroshells avoid the launch vehicle fairing diameter constraint by expanding to a larger diameter after the launch. Due to the potentially larger wetted area, deployable aeroshells achieve lower ballistic coefficients (well below 100 kg/m2), and if they are flown at shallower flight path angles, the entry vehicle can access trajectories with far lower decelerations (~50-60 g's), peak heat fluxes (~400 W/cm2) and peak pressures. The structural and TPS mass of the shallow-gamma, low-beta deployables are lower than their steep-gamma, high-beta rigid aeroshell counterparts at larger diameters, contributing to lower areal densities and potentially higher payload mass fractions. For example, at large diameters, deployables may attain aeroshell areal densities of 10 kg/m2 as opposed to 50 kg/m2 for rigid aeroshells. However, the low-beta, shallow-gamma paradigm also raises issues, such as the possibility of skip-out during entry. The shallow-gamma could also increase the landing footprint of the vehicle. Furthermore, the deployable entry systems may be flexible, so there could be fluid-structure interaction, especially in the high altitude, low-density regimes. The need for precision in guidance, navigation and control during entry also has to be better understood. This paper investigates some of the challenges facing the design of a shallow-gamma, low-beta entry system.

Dutta, Soumyo↗
NASA NTRS1992

Closed form solutions of constrained trajectories - Application in optimal ascent of aerospace plane

The present consideration of the flight trajectory of hypersonic aerospace vehicles subject to a class of path constraints notes the constrained dynamics to constitute a natural two-timescale system, so that problems of trajectory optimization and guidance can be dramatically simplified by means of the asymptotic analytical solutions thus obtained. An illustrative application in ascent trajectory optimization for an aerospace vehicle is presented.

Lu, Ping↗
NASA NTRSDate not supplied

Trajectory Operations of the Artemis I Mission

This paper describes the operational design and execution of the Artemis I trajectory. It was an operationally complex trajectory with powered lunar flybys and insertion into a Distant Retrograde Orbit (DRO). A joint team of trajectory analysts at the NASA Johnson Space Center (JSC) were responsible for the design and operation of nominal and off-nominal in-space trajectories. A process was developed to convert optimized reference trajectories into Orion burn plans that could be uplinked to the vehicle. During the mission, the joint flight controller and engineering team continuously evaluated upcoming translational burns using actual vehicle conditions, monitored the trajectory for opportunities to re-optimize the trajectory in order to reduce propellant usage, and prepared for potential off-nominal scenarios. Overall Orion in-space trajectory performance is compared tomission designs to demonstrate the success of the design and operations work-flows.

Artemis I↗
NASA NTRS2004

Space Shuttle Columbia Reconstruction Effort: An Analysis of the Thermal Protection System.

Upon the breakup of the Space Shuffle Columbia, which occurred during peak heating, at an altitude of 200,000 feet, and traveling at Mach 19, the daunting responsibility was placed upon National Aeronautics and Space Administration (NASA) to determine the root cause of the accident. One of the foremost tasks was to begin locating Columbia's debris, which was scattered acr ss a 2,000 square miles area of Texas and into western Louisiana. Each potential shuttle debris piece found was categorized and the longitude and latitude noted. This provided crucial data to the vehicle's breakup trajectory path and aided in locating key shuffle components. Since the airframe and tiles would be the primary focus during the reconstruction effort at the Kennedy Space center in Florida, packing and shipping was crucial to the preservation of the shape and condition of each recovered object. Over 83,000 pieces of Columbia, in a five-month timeframe, were located. The recovered tiles still bonded to structure indicated failures modes due to heating and aerodynamic loads. Also, different types of tile bond failure modes were observed, with adhesion failures at the Koropon layer or material failure at the tile's densification layer. Numerous tiles had the identification markings burned off, slumping of the Reaction-Cured Glass (RCG) coating, and of the Ll-900 silica base material. The conditions in which the tiles were found resulted in developing methods to aid in the identification of the tile. Correctly identifying each tile was critical to understanding the heating patterns of the tiles in relationship to the vehicle, since it was known that the lower left wing received the initial external tank foam impact. With the aid of multiple disciplines, the clues found in the TPS debris, lead to the root cause of the breakup of Columbia.

Micklos, Ann M.↗
NASA NTRS2019

Improvements to the Copernicus Trajectory Design and Optimization System for Complex Space Trajectories

The purpose of this assessment was to develop updates and new features for the NASA Copernicus Spacecraft Trajectory Design and Optimization analysis tool (version 5.0) for application to NASA programs and projects. These updates will significantly improve the ability to design and optimize complex trajectories over multiple trajectory phases; will allow the use of unique vehicle-specific guidance, control, and trajectory strategies and constraints; and the creation of an almost unlimited number of unique user-defined capabilities. The primary stakeholders for this assessment are the trajectory design and optimization analysts and engineers, and the chief engineers and project managers for existing programs, projects, and/or tasks that involve impulsive, finite burn, and/or continuous thrust trajectories (e.g., Sun, planet, comet, asteroid, halo orbit, Lagrange point, and distant retrograde orbit). The breadth of application spans the preliminary engineering and mission design concepts and optimization, to the development of candidate reference missions and integrated mission design for vehicle system design and operation, to the design and development of flight trajectories and associated propulsive maneuvers for real-time operations.

Murri, Daniel G.↗
NASA NTRS1973

Estimates of trapped radiation encountered on low-thrust trajectories through the Van Allen belts

Estimates were made of the number of trapped protons and electrons encountered by vehicles on low-thrust trajectories through the Van Allen belts. The estimates serve as a first step in assessing whether these radiations present a problem to on-board sensitive components and payload. The integrated proton spectra and electron spectra are presented for the case of a trajectory described by a vehicle with a constant-thrust acceleration A sub c equal to 0.001 meter/sq sec. This value of acceleration corresponds to a trip time of about 54 days from low earth orbit to synchronous orbit. It is shown that the time spent in the belts and hence the radiation encountered vary nearly inversely with the value of thrust acceleration. Thus, the integrated spectral values presented for the case of A sub c = 0.001 meter/sq sec can be generalized for any other value of thrust acceleration by multiplying them by the factor 0.001/A sub c.

Karp, I. M.↗