Search NASA⌕ Search

SEARCH · Search NASA

Results for “Injection Strategy”

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

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

At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

A multi-mission flyby strategy for the near-earth asteroids

Recent developments in miniaturization of spacecraft systems and science instruments have led to great interest in their application to deep space missions. A recent JPL study, for example, considered the design of a small spacecraft which could yield useful science data from fast flybys of the near-earth asteroids. Called AIM (Asteroid Investigation with Microspacecraft), three spacecrafts would be launched with Pegasus into low earth orbit (LEO), separated, and then remain there up to 2 weeks, each waiting for the opportune time to inject to a flyby with a preselected asteroid. Each spacecraft has its own kick stage for this injection maneuver. This paper briefly describes the spacecraft design and capability, the asteroid flyby opportunities available within the coming decade, and other possible mission scenarios which could take advantage of a single launch multi-spacecraft option.

Penzo, Paul A.↗
NASA NTRS2021

Orbit Determination for the Mars 2020 Mission

NASA’s Mars 2020 Mission successfully executed its Martian entry, descent, and landing sequence on February 18, 2021. Spacecraft navigation guided the spacecraft from its launch injection to entry into the Martian atmosphere; to do so, three trajectory correction maneuvers and two onboard state updates were performed. Navigation was required to provide an entry state with a maximum uncertainty of 2.8 km in position and 2.0 m/sec in velocity (3-σ); the flight performance was approximately 900 m and 66 cm/sec (3-σ). This paper presents the orbit determination models, measurements, filter strategies and performance throughout launch, cruise and Mars approach.

Wong, Mau↗
NASA NTRS2006

MOI to TEI : a Mars Sample Return strategy

This paper describes the issues and challenges related to the design of the rendezvous between the Earth Return Vehicle (ERV) and the Orbiting Sample (OS) for the Mars Sample Return (MSR) mission. In particular, attention will be focused on the strategy for 'optimizing' the intermediate segment of the rendezvous process, during which there are a great number of variables that must be considered and well understood.

rendezvous↗
NASA NTRS1992

Scramjet combustor and nozzle computations

The SHIP3D PNS code is used to conduct a 3D parametric study of a scramjet combustor and nozzle configuration at flight Mach numbers of 8, 10, 12, 16, and 20. The study demonstrated the powerful computational strategy of using a highly efficient, modified PNS code and a nonkinetic chemistry model to analyze test data and conduct parametric studies of scramjet combustors and nozzles. The mixing efficiency was found to decrease by 0.2 when the flight Mach number increased from 8 to 12 and showed little change at higher Mach numbers. The protruding ramp injector generated an oblique shock that impinged on the cowl. Ramp-slot injection produced better mixing, both because of the interaction of this shock with the slot fuel, and because of better fuel distribution in the duct. The former effect was enhanced when the ramp angle was increased. Mixing was also enhanced when the ramp fuel injection angle and sweep angle were increased.

Kamath, Pradeep S.↗
NASA NTRSDate not supplied

Optimized Trajectory Correction Burn Placement for the NASA Artemis II Mission

The NASA Artemis II mission represents the first time humans plan to return to the lunar vicinity in over 50 years with a crew traveling to the Moon in the Orion spacecraft on a free return trajectory. This first crewed mission of the Artemis program will evaluate human-rated elements of Orion in preparation to sending astronauts to the lunar surface. The selected free-return cislunar trajectory profile that is reminiscent of the Apollo 8 mission that nominally requires no additional translational burns following the trans-lunar injection (TLI) burn. Due to crew activity, maneuver execution errors, navigation uncertainty, orbit insertion errors, disturbance accelerations, and other system limitations; periodic trajectory corrections burns are necessary to ensure proper entry interface (EI) conditions are satisfied for a safe return to Earth. Robust trajectory optimization techniques are utilized to determine the optimized placements for the Artemis II trajectory correction burns that accounts for the crew schedule, both the primary and backup navigation systems, targeting strategies and burn plan configurations, spacecraft venting, thruster selection, and the integrated GN&C performance.

Linear Covariance Analysis↗
NASA NTRS2009

Launch Order, Launch Separation, and Loiter in the Constellation 1 1/2-Launch Solution

The NASA Constellation Program (CxP) is developing a two-element Earth-to-Orbit launch system to enable human exploration of the Moon. The first element, Ares I, is a human-rated system that consists of a first stage based on the Space Shuttle Program's solid rocket booster (SRB) and an upper stage that consists of a four-crew Orion capsule, a service module, and a Launch Escape System. The second element, Ares V, is a Saturn V-plus category launch system that consists of the core stage with a cluster of six RS-68B engines and augmented with two 5.5-segment SRBs, a Saturn-derived J-2X engine powering an Earth Departure Stage (EDS), and the lunar-lander vehicle payload, Altair. Initial plans called for the Ares V to be launched first, followed the next day by the Ares I. After the EDS performs the final portion of ascent and subsequent orbit circularization, the Orion spacecraft then performs a rendezvous and docks with the EDS and its Altair payload. Following checkout, the integrated stack loiters in low Earth orbit (LEO) until the appropriate Trans-Lunar Injection (TLI) window opportunity opens, at which time the EDS propels the integrated Orion Altair to the Moon. Successful completion of this 1 1/2-launch solution carries risks related to both the orbital lifetime of the assets and the probability of achieving the launch of the second vehicle within the orbital lifetime of the first. These risks, which are significant in terms of overall system design choices and probability of mission success, dictated a thorough reevaluation of the launch strategy, including the order of vehicle launch and the planned time period between launches. The goal of the effort described in this paper was to select a launch strategy that would result in the greatest possible expected system performance, while accounting for launch risks and the cost of increased orbital lifetime. Discrete Event Simulation (DES) model of the launch strategies was created to determine the probability of a second launch not occurring in a timely fashion (i.e., before the assets waiting in LEO expire). An overview of the launch strategy evaluation process is presented, along with results of specific cases that were analyzed. A high-level comparison of options is then presented, along with the conclusion derived from the analysis.

Stromgren, Chel↗
NASA NTRS2008

Demonstration of Active Combustion Control

The primary objective of this effort was to demonstrate active control of combustion instabilities in a direct-injection gas turbine combustor that accurately simulates engine operating conditions and reproduces an engine-type instability. This report documents the second phase of a two-phase effort. The first phase involved the analysis of an instability observed in a developmental aeroengine and the design of a single-nozzle test rig to replicate that phenomenon. This was successfully completed in 2001 and is documented in the Phase I report. This second phase was directed toward demonstration of active control strategies to mitigate this instability and thereby demonstrate the viability of active control for aircraft engine combustors. This involved development of high-speed actuator technology, testing and analysis of how the actuation system was integrated with the combustion system, control algorithm development, and demonstration testing in the single-nozzle test rig. A 30 percent reduction in the amplitude of the high-frequency (570 Hz) instability was achieved using actuation systems and control algorithms developed within this effort. Even larger reductions were shown with a low-frequency (270 Hz) instability. This represents a unique achievement in the development and practical demonstration of active combustion control systems for gas turbine applications.

Lovett, Jeffrey A.↗
NASA NTRS2015

Software Testbed for Developing and Evaluating Integrated Autonomous Subsystems

To implement fault tolerant autonomy in future space systems, it will be necessary to integrate planning, adaptive control, and state estimation subsystems. However, integrating these subsystems is difficult, time-consuming, and error-prone. This paper describes Intelliface/ADAPT, a software testbed that helps researchers develop and test alternative strategies for integrating planning, execution, and diagnosis subsystems more quickly and easily. The testbed's architecture, graphical data displays, and implementations of the integrated subsystems support easy plug and play of alternate components to support research and development in fault-tolerant control of autonomous vehicles and operations support systems. Intelliface/ADAPT controls NASA's Advanced Diagnostics and Prognostics Testbed (ADAPT), which comprises batteries, electrical loads (fans, pumps, and lights), relays, circuit breakers, invertors, and sensors. During plan execution, an experimentor can inject faults into the ADAPT testbed by tripping circuit breakers, changing fan speed settings, and closing valves to restrict fluid flow. The diagnostic subsystem, based on NASA's Hybrid Diagnosis Engine (HyDE), detects and isolates these faults to determine the new state of the plant, ADAPT. Intelliface/ADAPT then updates its model of the ADAPT system's resources and determines whether the current plan can be executed using the reduced resources. If not, the planning subsystem generates a new plan that reschedules tasks, reconfigures ADAPT, and reassigns the use of ADAPT resources as needed to work around the fault. The resource model, planning domain model, and planning goals are expressed using NASA's Action Notation Modeling Language (ANML). Parts of the ANML model are generated automatically, and other parts are constructed by hand using the Planning Model Integrated Development Environment, a visual Eclipse-based IDE that accelerates ANML model development. Because native ANML planners are currently under development and not yet sufficiently capable, the ANML model is translated into the New Domain Definition Language (NDDL) and sent to NASA's EUROPA planning system for plan generation. The adaptive controller executes the new plan, using augmented, hierarchical finite state machines to select and sequence actions based on the state of the ADAPT system. Real-time sensor data, commands, and plans are displayed in information-dense arrays of timelines and graphs that zoom and scroll in unison. A dynamic schematic display uses color to show the real-time fault state and utilization of the system components and resources. An execution manager coordinates the activities of the other subsystems. The subsystems are integrated using the Internet Communications Engine (ICE). an object-oriented toolkit for building distributed applications.

Testbed↗
NASA NTRSDate not supplied

Synthetic Failure Mode Generation for Resilience Analysis and Failure Mechanism Discovery

Traditional risk-based design processes seek to mitigate operational hazards by manually identifying possible faults and corresponding mitigation strategies—a tedious process which critically relies on the designer’s limited knowledge. Resilience-based design, on the other hand, seeks to embody generic hazard-mitigating properties in the system to mitigate unknown hazards, often by modelling the system's response to potential hazardous events. This work adapts this approach to the traditional risk-based design process to synthetically generate hazardous modes, by representing them as a unique combination of internal component health-states which can then be injected and simulated in a model of the system failure dynamics. The design process may then reduce the risk of unknown internal hazards by iteratively mitigating the effects of these modes. The performance of this approach is evaluated in a model of an autonomous rover, where cluster analysis shows that elaborating the space of synthetic faults in the drive system using this approach uncovers a wider range of possible hazardous trajectories and failure consequences within each trajectory. However, this increase in hazard information comes at a high computational expense, highlighting the need for advanced, efficient methods to search and sample the hazard space.

Simulation↗
NASA NTRSDate not supplied

Synthetic Failure Mode Generation for Resilience Analysis and Failure Mechanism Discovery

Traditional risk-based design processes seek to mitigate operational hazards by manually identifying possible faults and corresponding mitigation strategies—a tedious process which critically relies on the designer’s limited knowledge. Resilience-based design, on the other hand, seeks to embody generic hazard-mitigating properties in the system to mitigate unknown hazards, often by modelling the system's response to potential hazardous events. This work adapts this approach to the traditional risk-based design process to synthetically generate hazardous modes, by representing them as a unique combination of internal component health-states which can then be injected and simulated in a model of the system failure dynamics. The design process may then reduce the risk of unknown internal hazards by iteratively mitigating the effects of these modes. The performance of this approach is evaluated in a model of an autonomous rover, where cluster analysis shows that elaborating the space of synthetic faults in the drive system using this approach uncovers a wider range of possible hazardous trajectories and failure consequences within each trajectory. However, this increase in hazard information comes at a high computational expense, highlighting the need for advanced, efficient methods to search and sample the hazard space.

Simulation↗
NASA NTRS1998

On the Generation and Use of TCP Acknowledgments

This paper presents a simulation study of various TCP acknowledgment generation and utilization techniques. We investigate the standard version of TCP and the two standard acknowledgment strategies employed by receivers: those that acknowledge each incoming segment and those that implement delayed acknowledgments. We show the delayed acknowledgment mechanism hurts TCP performance, especially during slow start. Next we examine three alternate mechanisms for generating and using acknowledgments designed to mitigate the negative impact of delayed acknowledgments. The first method is to generate delayed ACKs only when the sender is not using the slow start algorithm. The second mechanism, called byte counting, allows TCP senders to increase the amount of data being injected into the network based on the amount of data acknowledged rather than on the number of acknowledgments received. The last mechanism is a limited form of byte counting. Each of these mechanisms is evaluated in a simulated network with no competing traffic, as well as a dynamic environment with a varying amount of competing traffic. We study the costs and benefits of the alternate mechanisms when compared to the standard algorithm with delayed ACKs.

Allman, Mark↗
NASA NTRSDate not supplied

Turbine Electrified Energy Management with Model Predictive Control

Affordability, sustainability, and efficiency are primary motivators driving the future of NASA aeronautics research. These factors are realized, in part, through the development and implementation of new technologies and strategies enabling efficient, affordable, and safe hybrid-electric aircraft. Research supporting electrified aircraft propulsion control systems exemplifies such new methodologies, offering varied opportunities to integrate electric machines with gas-based turbine engines. For hybrid-electric propulsion systems, current conceptual architectures seek to introduce energy storage and exploit electrical power system components to assist gas-based system components. Capitalizing on the electric machines in hybridized engines, Turbine Electrified Energy Management (TEEM) is a control approach that enhances transient operability to improve overall propulsion and vehicle efficiency by injecting or extracting power from engine shafts. Traditionally implemented with proportional-integral (PI) control, this study expands the application of TEEM by presenting model predictive control (MPC) schemes to execute the TEEM concept. Via cost function design and constraint selection, the transient operability goals for TEEM are considered in the controller designs. The proposed MPCs are simulated on a nonlinear turbofan engine model at two environmental conditions, with comparisons drawn to a baseline PI. Performance is evaluated using compressor maps and two TEEM-specific metrics: transient stack usage and transient excursion integral. Simulation results reveal the developed schemes perform comparably to the benchmark controller and can be implemented in two distinct configurations. Potential modifications for future investigations include cost function measures that optimize energy use, additional performance effectiveness measures, and battery storage capabilities.

Elyse D. Hill↗
NASA NTRS1996

Orbital Injection of the SEDSAT Satellite: Tethered Systems Dynamics and Flight Data Analysis

This report deals with the following topics which are all related to the orbital injection of the SEDSAT satellite: Dynamics and Stability of Tether Oscillations after the First Cut. The dynamics of the tether after the first cut (i.e., without the Shuttle attached to it) is investigated. The tether oscillations with the free end are analyzed in order to assess the stability of the rectilinear configuration in between the two tether cuts; analysis of Unstable Modes. The unstable modes that appear for high libration angles are further investigated in order to determine their occurrences and the possible transition from bound librations to rotations; Orbital Release Strategies for SEDSAT. A parametric analysis of the orbital decay rate of the SEDSAT satellite after the two tether cuts has been carried out as a function of the following free parameters: libration amplitude at the end of deployment, deviation angle from LV at the first cut, and orbital anomaly at the second cut. The values of these parameters that provide a minimum orbital decay rate of the satellite (after the two cuts) have been computed; and Dynamics and Control of SEDSAT. The deployment control law has been modified to cope with the new ejection velocity of the satellite from the Shuttle cargo bay. New reference profiles have been derived as well as new control parameters. Timing errors at the satellite release as a function of the variations of the initial conditions and the tension model parameters have been estimated for the modified control law.

Lorenzini, Enrico C.↗
NASA NTRS2010

NASA's Soil Moisture Active and Passive (SMAP) Mission

The Soil Moisture Active-Passive (SMAP) Mission is one of the first Earth observation satellites being formulated by NASA in response to the 2007 National Research Council s Decadal Survey. SMAP will make global measurements of soil moisture at the Earth's land surface and its freeze-thaw state. These measurements will allow significantly improved estimates of water, energy and carbon transfers between the land and atmosphere. Soil moisture measurements are also of great importance in assessing flooding and monitoring drought. Knowledge gained from SMAP observations can help mitigate these natural hazards, resulting in potentially great economic and social benefits. SMAP observations of soil moisture and freeze/thaw timing over the boreal latitudes will also reduce a major uncertainty in quantifying the global carbon balance and help to resolve an apparent missing carbon sink over land. The SMAP mission concept will utilize an L-band radar and radiometer sharing a rotating 6-meter mesh reflector antenna flying in a 680 km polar orbit with an 8-day exact ground track repeat aboard a 3-axis stabilized spacecraft to provide high-resolution and high-accuracy global maps of soil moisture and freeze/thaw state every two to three days. In addition, the SMAP project will use these surface observations with advanced modeling and data assimilation to provide estimates of deeper root-zone soil moisture and net ecosystem exchange of carbon. SMAP recently completed its Phase A Mission Concept Study Phase for NASA and transitioned into Phase B (Formulation and Detailed Design). A number of significant accomplishments occurred during this initial phase of mission development. The SMAP project held several open meetings to solicit community feedback on possible science algorithms, prepared preliminary draft Algorithm Theoretical Basis Documents (ATBDs) for each mission science product, and established a prototype algorithm testbed to enable testing and evaluation of the performance of candidate algorithms. SMAP conducted an Applications Workshop in September 2009 to coordinate with potential application users interested in the mission data. A draft Applications Plan describing the Project s planned outreach to potential applications users has been prepared and will be updated during Phase B. SMAP made a significant evaluation of the potential terrestrial radio frequency interference (RFI) source environment and established radiometer and radar flight hardware and ground processing mitigation approaches. SMAP finalized its science orbit and orbit injection approach to optimize launch mass and prepared launch and commissioning scenarios and timeline. A science data communications approach was developed to maximize available science data volume to improve science margins while maintaining moderately short data product latencies to support many potential applications using existing ground assets and with minimum impact to the flight system. SMAP developed rigid multi-body and flexible body dynamics and control models and system designs for the 6-meter rotating instrument reflector-boom assembly (RBA) and flight system to confirm pointing and control performance, and devised strategies to efficiently implement on-orbit balancing if needed. Industry partners were selected for the spin mechanism assembly (SMA) and RBA. Preliminary designs for the radar and radiometer were initiated, including constructing breadboards of key assemblies.

Kellogg, Kent↗
NASA NTRS2011

Mars Sample Return - Launch and Detection Strategies for Orbital Rendezvous

This study sets forth conceptual mission design strategies for the ascent and rendezvous phase of the proposed NASA/ESA joint Mars Sample Return Campaign. The current notional mission architecture calls for the launch of an acquisition/cache rover in 2018, an orbiter with an Earth return vehicle in 2022, and a fetch rover and ascent vehicle in 2024. Strategies are presented to launch the sample into a coplanar orbit with the Orbiter which facilitate robust optical detection, orbit determination, and rendezvous. Repeating ground track orbits exist at 457 and 572 km which provide multiple launch opportunities with similar geometries for detection and rendezvous.

Orbiter's trans-Earth injection (TEI):↗
NASA NTRS2014

2016 Mars Insight Mission Design and Navigation

Scheduled for a launch in the 2016 Earth to Mars opportunity, the Interior Exploration using Seismic Investigations, Geodesy, and Heat Transport (InSight) Mission will arrive to Mars in late September 2016 with the primary objective of placing a science lander on the surface of the Red Planet followed by the deployment of two science instruments to investigate the fundamental processes of terrestrial planet formation and evolution. In order to achieve a successful landing, the InSight Project has selected a launch/arrival strategy that satisfies the following key and driving requirements: (1) Deliver a total launch mass of 727 kg, (2) target a nominal landing site with a cumulative Delta V99 less than 30 m/s, and (3) approach EDL with a V-infinity upper limit of 3.941 km/s and (4) an entry flight-path angle (EFPA) of -12.5 +/- 0.26 deg, 3-sigma; the InSight trajectories have been designed such that they (5) provide UHF-band communications via Direct-To-Earth and MRO from Entry through landing plus 60 s, (6) with injection aimpoints biased away from Mars such that the probability of the launch vehicle upper stage impacting Mars is less than 1.0 X 10(exp 4) for fifty years after launch, and (7) non-nominal impact probabilities due to failure during the Cruise phase less than 1.0 X 10(exp 2).

Landers↗
NASA NTRS1981

Development of a low NO/x/ lean premixed annular combustor

The results of an experimental atmospheric rig test program developed to define a low NO(x) annular combustor of 0.66 m diameter for high-altitude aircraft applications are presented. The test program strategy adopted evaluates the emission characteristics of a baseline configuration and subsequently examines the sensitivity of the emission signatures to variations in the key design features. The lean premixed combustor in axisymmetric annular form demonstrates the capability of operating at reduced-pressure, simulated high-altitude, supersonic cruise conditions with NO(x) emissions below 1.0 g NO2/kg fuel. The testing shows that for the full range of low emissions operation from idle to cruise, a variable dilution port system is necessary, but that fuel-switching can be avoided and a single fuel injection system used.

Roberts, P. B.↗
NASA NTRS2016

Innovative Approach for Developing Spacecraft Interior Acoustic Requirement Allocation

The Orion Multi-Purpose Crew Vehicle (MPCV) is an American spacecraft for carrying four astronauts during deep space missions. This paper describes an innovative application of Power Injection Method (PIM) for allocating Orion cabin continuous noise Sound Pressure Level (SPL) limits to the sound power level (PWL) limits of major noise sources in the Environmental Control and Life Support System (ECLSS) during all mission phases. PIM is simulated using both Statistical Energy Analysis (SEA) and Hybrid Statistical Energy Analysis-Finite Element (SEA-FE) models of the Orion MPCV to obtain the transfer matrix from the PWL of the noise sources to the acoustic energies of the receivers, i.e., the cavities associated with the cabin habitable volume. The goal of the allocation strategy is to control the total energy of cabin habitable volume for maintaining the required SPL limits. Simulations are used to demonstrate that applying the allocated PWLs to the noise sources in the models indeed reproduces the SPL limits in the habitable volume. The effects of Noise Control Treatment (NCT) on allocated noise source PWLs are investigated. The measurement of source PWLs of involved fan and pump development units are also discussed as it is related to some case-specific details of the allocation strategy discussed here.

Chu, S. Reynold↗