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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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NASA NTRS1992

Interplanetary program to optimize simulated trajectories (IPOST). Volume 4: Sample cases

The Interplanetary Program to Optimize Simulated Trajectories (IPOST) is intended to support many analysis phases, from early interplanetary feasibility studies through spacecraft development and operations. The IPOST output provides information for sizing and understanding mission impacts related to propulsion, guidance, communications, sensor/actuators, payload, and other dynamic and geometric environments. IPOST models three degree of freedom trajectory events, such as launch/ascent, orbital coast, propulsive maneuvering (impulsive and finite burn), gravity assist, and atmospheric entry. Trajectory propagation is performed using a choice of Cowell, Encke, Multiconic, Onestep, or Conic methods. The user identifies a desired sequence of trajectory events, and selects which parameters are independent (controls) and dependent (targets), as well as other constraints and the cost function. Targeting and optimization are performed using the Standard NPSOL algorithm. The IPOST structure allows sub-problems within a master optimization problem to aid in the general constrained parameter optimization solution. An alternate optimization method uses implicit simulation and collocation techniques.

Hong, P. E.↗
NASA NTRS1992

Interplanetary Program to Optimize Simulated Trajectories (IPOST). Volume 2: Analytic manual

The Interplanetary Program to Optimize Space Trajectories (IPOST) is intended to support many analysis phases, from early interplanetary feasibility studies through spacecraft development and operations. The IPOST output provides information for sizing and understanding mission impacts related to propulsion, guidance, communications, sensor/actuators, payload, and other dynamic and geometric environments. IPOST models three degree of freedom trajectory events, such as launch/ascent, orbital coast, propulsive maneuvering (impulsive and finite burn), gravity assist, and atmospheric entry. Trajectory propagation is performed using a choice of Cowell, Encke, Multiconic, Onestep, or Conic methods. The user identifies a desired sequence of trajectory events, and selects which parameters are independent (controls) and dependent (targets), as well as other constraints and the cost function. Targeting and optimization is performed using the Stanford NPSOL algorithm. IPOST structure allows subproblems within a master optimization problem to aid in the general constrained parameter optimization solution. An alternate optimization method uses implicit simulation and collocation techniques.

Hong, P. E.↗
NASA NTRS1998

The Control System for the X-33 Linear Aerospike Engine

The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.

Jackson, Jerry E.↗
NASA NTRS1998

Ascent, Transition, Entry, and Abort Guidance Algorithm Design for the X-33 Vehicle

One of the primary requirements for X-33 is that it be capable of flying autonomously. That is, onboard computers must be capable of commanding the entire flight from launch to landing, including cases where a single engine failure abort occurs. Guidance algorithms meeting these requirements have been tested in simulation and have been coded into prototype flight software. These algorithms must be sufficiently robust to account for vehicle and environmental dispersions, and must issue commands that result in the vehicle operating, within all constraints. Continual tests of these algorithms (and modifications as necessary) will occur over the next year as the X-33 nears its first flight. This paper describes the algorithms in use for X-33 ascent, transition, and entry flight, as well as for the powered phase of PowerPack-out (PPO) aborts (equivalent in thrust impact to losing an engine). All following discussion refers to these phases of flight when discussing guidance. The paper includes some trajectory results and results of dispersion analysis.

Hanson, John M.↗
NASA NTRS1998

Safety Analysis of FMS/CTAS Interactions During Aircraft Arrivals

This grant funded research on human-computer interaction design and analysis techniques, using future ATC environments as a testbed. The basic approach was to model the nominal behavior of both the automated and human procedures and then to apply safety analysis techniques to these models. Our previous modeling language, RSML, had been used to specify the system requirements for TCAS II for the FAA. Using the lessons learned from this experience, we designed a new modeling language that (among other things) incorporates features to assist in designing less error-prone human-computer interactions and interfaces and in detecting potential HCI problems, such as mode confusion. The new language, SpecTRM-RL, uses "intent" abstractions, based on Rasmussen's abstraction hierarchy, and includes both informal (English and graphical) specifications and formal, executable models for specifying various aspects of the system. One of the goals for our language was to highlight the system modes and mode changes to assist in identifying the potential for mode confusion. Three published papers resulted from this research. The first builds on the work of Degani on mode confusion to identify aspects of the system design that could lead to potential hazards. We defined and modeled modes differently than Degani and also defined design criteria for SpecTRM-RL models. Our design criteria include the Degani criteria but extend them to include more potential problems. In a second paper, Leveson and Palmer showed how the criteria for indirect mode transitions could be applied to a mode confusion problem found in several ASRS reports for the MD-88. In addition, we defined a visual task modeling language that can be used by system designers to model human-computer interaction. The visual models can be translated into SpecTRM-RL models, and then the SpecTRM-RL suite of analysis tools can be used to perform formal and informal safety analyses on the task model in isolation or integrated with the rest of the modeled system. We had hoped to be able to apply these modeling languages and analysis tools to a TAP air/ground trajectory negotiation scenario, but the development of the tools took more time than we anticipated.

Nancy G. Leveson↗
NASA NTRS1999

Departure Energies, Trip Times and Entry Speeds for Human Mars Missions

The study examines how the mission design variables departure energy, entry speed, and trip time vary for round-trip conjunction-class Mars missions. These three parameters must be balanced in order to produce a mission that is acceptable in terms of mass, cost, and risk. For the analysis, a simple, massless-planet trajectory program was employed. The premise of this work is that if the trans-Mars and trans-Earth injection stages are designed for the most stringent opportunity in the energy cycle, then there is extra energy capability in the "easier" opportunities which can be used to decrease the planetary entry speed, or shorten the trip time. Both of these effects are desirable for a human exploration program.

Munk, Michelle M.↗
NASA NTRS2000

Postural control model interpretation of stabilogram diffusion analysis

Collins and De Luca [Collins JJ. De Luca CJ (1993) Exp Brain Res 95: 308-318] introduced a new method known as stabilogram diffusion analysis that provides a quantitative statistical measure of the apparently random variations of center-of-pressure (COP) trajectories recorded during quiet upright stance in humans. This analysis generates a stabilogram diffusion function (SDF) that summarizes the mean square COP displacement as a function of the time interval between COP comparisons. SDFs have a characteristic two-part form that suggests the presence of two different control regimes: a short-term open-loop control behavior and a longer-term closed-loop behavior. This paper demonstrates that a very simple closed-loop control model of upright stance can generate realistic SDFs. The model consists of an inverted pendulum body with torque applied at the ankle joint. This torque includes a random disturbance torque and a control torque. The control torque is a function of the deviation (error signal) between the desired upright body position and the actual body position, and is generated in proportion to the error signal, the derivative of the error signal, and the integral of the error signal [i.e. a proportional, integral and derivative (PID) neural controller]. The control torque is applied with a time delay representing conduction, processing, and muscle activation delays. Variations in the PID parameters and the time delay generate variations in SDFs that mimic real experimental SDFs. This model analysis allows one to interpret experimentally observed changes in SDFs in terms of variations in neural controller and time delay parameters rather than in terms of open-loop versus closed-loop behavior.

NASA Program Biomedical Research and Countermeasur↗
NASA NTRS2004

Status of Low Thrust Work at JSC

High performance low thrust (solar electric, nuclear electric, variable specific impulse magnetoplasma rocket) propulsion offers a significant benefit to NASA missions beyond low Earth orbit. As NASA (e.g., Prometheus Project) endeavors to develop these propulsion systems and associated power supplies, it becomes necessary to develop a refined trajectory design capability that will allow engineers to develop future robotic and human mission designs that take advantage of this new technology. This ongoing work addresses development of a trajectory design and optimization tool for assessing low thrust (and other types) trajectories. This work targets to advance the state of the art, enable future NASA missions, enable science drivers, and enhance education. This presentation provides a summary of the low thrust-related JSC activities under the ISP program and specifically, provides a look at a new release of a multi-gravity, multispacecraft trajectory optimization tool (Copernicus) along with analysis performed using this tool over the past year.

Condon, Gerald L.↗
NASA NTRS2011

STS-134 Re-Rendezvous Design History

In preparation to provide the capability for the Orion spacecraft to rendezvous with the International Space Station (ISS), a new suite of relative navigation sensors are in development and will be tested on one of the final Space Shuttle missions to ISS. The National Aeronautics and Space Administration (NASA) commissioned a flight test of prototypes of the instruments on STS-134, in order to test their performance in the space environment during the nominal rendezvous and docking, as well as a re-rendezvous dedicated to testing the prototype sensors following the undocking of the Space Shuttle Orbiter at the end of the mission. Unlike the initial rendezvous and docking, the re-rendezvous profile would replicate the newly designed Orion coelliptic approach trajectory, something never before attempted with the Shuttle Orbiter. Therefore, there were a number of new parameters that needed to be conceived of, designed, and tested for this re-rendezvous to make the flight test successful. And all of this work had to be integrated with the normal operations of the ISS and Shuttle and had to conform to the constraints of the mission and vehicles. The result of this work is a separation and re-rendezvous trajectory design that will prove not only the design of the relative navigation sensors for the Orion vehicle, but also will serve as a proof of concept for the Orion rendezvous trajectory itself. This document presents the analysis and decision making process involved in attaining the final STS-134 re-rendezvous design.

Stuit, Timothy D.↗
NASA NTRS2011

Designing the STS-134 Re-Rendezvous: A Preparation for Future Crewed Rendezvous Missions

In preparation to provide the capability for the Orion spacecraft, also known as the Multi-Purpose Crew Vehicle (MPCV), to rendezvous with the International Space Station (ISS) and future spacecraft, a new suite of relative navigation sensors are in development and were tested on one of the final Space Shuttle missions to ISS. The National Aeronautics and Space Administration (NASA) commissioned a flight test of prototypes of the Orion relative navigation sensors on STS-134, in order to test their performance in the space environment during the nominal rendezvous and docking, as well as a re-rendezvous dedicated to testing the prototype sensors following the undocking of the Space Shuttle orbiter at the end of the mission. Unlike the rendezvous and docking at the beginning of the mission, the re-rendezvous profile replicates the newly designed Orion coelliptic approach trajectory, something never before attempted with the shuttle orbiter. Therefore, there were a number of new parameters that needed to be conceived of, designed, and tested for this rerendezvous to make the flight test successful. Additionally, all of this work had to be integrated with the normal operations of the ISS and shuttle and had to conform to the constraints of the mission and vehicles. The result of this work is a separation and rerendezvous trajectory design that would not only prove the design of the relative navigation sensors for the Orion vehicle, but also would serve as a proof of concept for the Orion rendezvous trajectory itself. This document presents the analysis and decision making process involved in attaining the final STS-134 re-rendezvous design.

Stuit, Timothy D.↗
NASA NTRS2011

Modeling Off-Nominal Recovery in NextGen Terminal-Area Operations

Robust schedule-based arrival management requires efficient recovery from off-nominal situations. This paper presents research on modeling off-nominal situations and plans for recovering from them using TRAC, a route/airspace design, fast-time simulation, and analysis tool for studying NextGen trajectory-based operations. The paper provides an overview of a schedule-based arrival-management concept and supporting controller tools, then describes TRAC implementations of methods for constructing off-nominal scenarios, generating trajectory options to meet scheduling constraints, and automatically producing recovery plans.

Callantine, Todd J.↗
NASA NTRS2014

The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30 to 420 days in duration and assuming chemical propulsion. Similar studies have been reported assuming high power electric propulsion for manned NEO rendezvous missions (11). The delta V requirement breakdown and mission profile data from references 10 and 11 are used as a basis for sizing the NEO Scout spacecraft and for conducting preliminary feasibility assessments using the Tsiokolvsky rocket equation, a (worst-case) delta V requirement of 10 km/sec, and a maximum spacecraft dry mass of 20 kg. Using chemical propellant for a 10 km/sec delta V drives spacecraft wet mass well above 300 kg so that chemical propulsion is a non-starter for the proposed mission profile and spacecraft wet mass limits. In contrast, a solar electric propulsion system needs only 8 kg of Xe propellant to accelerate the spacecraft to 10 km/sec in 163 days with 0.02 N of thrust and 500 W of power from1.6 sq m of 29% efficient solar panels. In a second example, accelerating a 4 kg payload to 7 km/sec over 180 days requires about 6.7 kg of propellant and 1.2 kg of solar panels (12 kg total spacecraft wet mass).

Koontz, Steven L.↗
NASA NTRS2018

Using Concurrent Cardiovascular Information to Augment Survival Time Data for Evaluating Orthostatic Tilt Test Performance

Head-up tilt (HUT) tests often are used in research to measure orthostatic intolerance (OI) (inability to appropriately control blood pressure while upright) in clinical populations and otherwise healthy individuals after interventions. Post-space flight orthostatic intolerance is a well-known phenomenon, and countermeasures to its development has been an active area of research at NASA. In the NASA HUT protocol, subjects lie horizontally on an automatic tilt table for baseline measurements before being raised to 80deg head-up tilt for a defined period of time or until signs or symptoms of presyncope ensues (light-headedness, nausea, dizziness, sweating, weakness or fainting). Multiple measures are collected to evaluate the cardiovascular system's ability to respond appropriately to the orthostatic challenge. However if the intended duration of the HUT is short, the ability to detect changes in OI due to an intervention or its prevention by a countermeasure may be limited by a small number of failures to permit comparisons based on survival time alone. Thus, the time-trajectory of the cardiovascular data becomes an important additional source of information. In particular, we will show how various measures of trajectory variability can effectively augment survival analysis for the assessment of OI in a joint model when high censoring rates are present.

Feiveson, Alan H.↗
NASA NTRS2018

MMS Examination of FTEs at the Earth's Subsolar Magnetopause

Determining the magnetic field structure, electric currents, and plasma distributions within flux transfer event (FTE)-type flux ropes is critical to the understanding of their origin, evolution, and dynamics. Here the Magnetospheric Multiscale mission's high-resolution magnetic field and plasma measurements are used to identify FTEs in the vicinity of the subsolar magnetopause. The constant-alpha flux rope model is used to identify quasi-force free flux ropes and to infer the size, the core magnetic field strength, the magnetic flux content, and the spacecraft trajectories through these structures. Our statistical analysis determines a mean diameter of 1,700 ± 400 km (~30 ± 9 d(sub i)) and an average magnetic flux content of 100 ± 30 kWb for the quasi-force free FTEs at the Earth's subsolar magnetopause which are smaller than values reported by Cluster at high latitudes. These observed nonlinear size and magnetic flux content distributions of FTEs appear consistent with the plasmoid instability theory, which relies on the merging of neighboring, small-scale FTEs to generate larger structures. The ratio of the perpendicular to parallel components of current density, R(sub J), indicates that our FTEs are magnetically force-free, defined as R(sub J) < 1, in their core regions (<0.6 R(sub flux rope)). Plasma density is shown to be larger in smaller, newly formed FTEs and dropping with increasing FTE size. It is also shown that parallel ion velocity dominates inside FTEs with largest plasma density. Field-aligned flow facilitates the evacuation of plasma inside newly formed FTEs, while their core magnetic field strengthens with increasing FTE size.

Akhavan-Tafti, M.↗
NASA NTRSDate not supplied

Artemis I Trajectory Design and Optimization

This paper presents the overall trajectory design and optimization process for NASA’s Artemis I mission to send an uncrewed Orion vehicle to a lunar Distant Retrograde Orbit (DRO). The on-orbit trajectory begins at the Space Launch System (SLS) core separation and ends at the Orion service module Earth Entry Interface (EI) point. The details of the trajectory optimization process are presented, including design of nominal and extended mission options, launch windows, and abort options. Novel design techniques are also discussed to account for contingencies, such as using auxiliary thrusters to protect against main engine failure and applying trajectory shaping to mitigate or reduce eclipse durations.

Artemis I↗
NASA NTRSDate not supplied

Linear Covariance Techniques to Analyze a Multi-Vehicle, Multi-GN&C System with Applications to Rendezvous in a Near Rectilinear Halo Orbit

Typically for a rendezvous and docking scenario, only a single vehicle is considered the active vehicle. As a result, the target vehicle’s GN&C system is typically not a factor in the integrated performance analysis as it assumes a passive role. However, for upcoming NASA Artemis missions the active vehicle is both the chaser and target spacecraft. In addition, both vehicles are cooperating with one another, sharing telemetry data, and relying on data extracted from the GN&C system of the other. The process also includes a team of ground support personnel in mission control are tracking and monitoring each spacecraft uplinking state estimates and targeting solutions to support mission operations and enhance the onboard flight system performance. Being able to quickly analyze the impact of both vehicles with two different GN&C systems that also interact with the ground that does its own navigation and targeting uploads is critical. This paper outlines how to perform this rapid analysis using linear covariance techniques and applies them to a rendezvous scenario initiated in low lunar orbit and completed in a Near Rectilinear Halo Orbit (NRHO) representative of the NASA Artemis III mission.

Linear Covariance Analysis↗
NASA NTRSDate not supplied

Ballistic Lunar Return Trajectories for Sustainable Cargo Return and Entry System Technology Development

As part of the sustained approach for the lunar Gateway and Artemis program as a whole, NASA is extending the logistics supply chain beyond low Earth orbit and to the Moon. This supply chain includes the possibility of lunar sample and cargo return. To enable these future return missions, the possibility for incorporating demonstration payloads including various entry, descent, and landing technologies is examined. Possible implementations include deployable entry vehicles, high speed sample return capsules, aeroassist technologies, and control technologies for guided hypersonic flight. The mission concepts utilize the secondary payload capabilities provided on a relatively low-cost logistic module. The logistic module may perform close Earth flyby, pointing, and release of return systems prior to disposal, with reentry velocities for payloads on the order of 11.5 km/sec. In this paper, we study the various return mission architectures available during the Artemis program to define the range of mission possibilities. Potential options include a reusable logistics module using a hypersonic inflatable aerodynamic decelerator, externally mounted entry system vehicle technologies, and a reusable sample return vehicle using a similar inflatable architecture with a feathered configuration with applied aerodynamic control. In these cases, the use of a ballistic lunar return trajectory is assumed, and a sensitivity analysis of midcourse corrections and the possibility of a lunar gravity assist for refining Earth entry interface points is provided. Furthermore, a novel controller for controlling a vehicle during reentry imposing heating limitations is introduced.

Matthew M. Wittal↗