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NASA NTRS2010

Orion Entry Display Feeder and Interactions with the Entry Monitor System

The Orion spacecraft is designed to return astronauts to a landing within 10 km of the intended landing target from low Earth orbit, lunar direct-entry, and lunar skip-entry trajectories. Al pile the landing is nominally controlled autonomously, the crew can fly precision entries manually in the event of an anomaly. The onboard entry displays will be used by the crew to monitor and manually fly the entry, descent, and landing, while the Entry Monitor System (EMS) will be used to monitor the health and status of the onboard guidance and the trajectory. The entry displays are driven by the entry display feeder, part of the Entry Monitor System (EMS). The entry re-targeting module, also part of the EMS, provides all the data required to generate the capability footprint of the vehicle at any point in the trajectory, which is shown on the Primary Flight Display (PFD). It also provides caution and warning data and recommends the safest possible re-designated landing site when the nominal landing site is no longer within the capability of the vehicle. The PFD and the EMS allow the crew to manually fly an entry trajectory profile from entry interface until parachute deploy having the flexibility to manually steer the vehicle to a selected landing site that best satisfies the priorities of the crew. The entry display feeder provides data from the ENIS and other components of the GNC flight software to the displays at the proper rate and in the proper units. It also performs calculations that are specific to the entry displays and which are not made in any other component of the flight software. In some instances, it performs calculations identical to those performed by the onboard primary guidance algorithm to protect against a guidance system failure. These functions and the interactions between the entry display feeder and the other components of the EMS are described.

Baird, Darren↗

NASA NTRS1984

Analytical characterization of AOTV perigee aerothermodynamic regime

In preliminary design studies of AOTV (Aeroassisted Orbital Transfer Vehicle) configurations, it is desirable to have a simple analytical method for characterizing the aerothermodynamic regime at skip trajectory perigee as a function of vehicle L/D and m/CDA (ballistic coefficient). The present study derives an approximate perigee solution from the generalized equations of motion. For a prescribed entry velocity vector, the solution determines a Zeta function (proportional to density divided by m/CDA) at perigee as a parametric function of the exit velocity vector and (constant) L/D. The perigee Zeta function then determines perigee density (or altitude) as a parametric function of m/CDA. The solution allows the following classic aerothermodynamic parameters to be determined at perigee as parametric functions of m/CDA, L/D, and the exit velocity vector: Reynolds Number (viscous effects), Mach Number (compressibility effects), Knudsen Number (rarefaction effects), sphere and disk bow shock standoff distance, Damkohler Number (relaxation effects), viscous correlation parameter (viscid-inviscid interactions), and Stanton Number (convective heat transfer). Results of the analysis are given for low L/D and mid L/D AOTV configurations on return from geosynchronous and L5 orbits. It is concluded the method successfully provides preliminary estimates of the aerothermodynamic parameters through the use of simple algebraic equations and plots.

Desautel, D.↗

NASA NTRS1998

Arcjet Semi-Elliptic Nozzle Simulations and Validation in Support of TPS Testing

Much of the ground based testing of advanced thermal protection system (TPS) components for the X33 program is done in arc-heated wind tunnels such as those located in the Arc-Jet Complex at NASA Ames Research Center. These facilities are capable of simulating the high temperature, chemically reacting flow environment experienced by the vehicle during flight. This allows one to test critical design issues such as maximum reuse temperatures, seals, gaps, and increases in heating due to interfaces between different materials. Computational fluid dynamics (CFD) has evolved to the point where it now can be used in the vehicle design process for accurate and timely prediction of trajectory based aerothermal heating environments for re-entry vehicles. It can also be used for simulation of the flow environments in ground based facilities such as arcjets. By utilization of the same CFD code and solution methodology, the important differences between ground test and flight may be quantified. The goal of this paper is to utilize CFD to provide validated simulations of the flow environment in the NASA-Ames semi elliptic nozzle arcjet facilities. The validation of the ground simulations will come From comparison to existing calibration data. Specific tests in support of the X33 TPS test program will ilso be simulated. In this manner, the differences between the ground test simulation and the flight environment can be identified for a measure of ground test to flight traceability.

Loomis, Mark↗

NASA NTRS2016

Mission Design Considerations for Mars Cargo of the Human Spaceflight Architecture Team's Evolvable Mars Campaign

Recent NASA interest in human missions to Mars has led to an Evolvable Mars Campaign by the agency's Human Architecture Team. Delivering the crew return propulsion stages and Mars surface landers, SEP based systems are employed because of their high specific impulse characteristics enabling missions requiring less propellant although with longer transfer times. The Earth departure trajectories start from an SLS launch vehicle delivery orbit and are spiral shaped because of the low SEP thrust. Previous studies have led to interest in assessing the divide in trip time between the Earth departure and interplanetary legs of the mission for a representative SEP cargo vehicle.

mars missions↗

NASA NTRS1972

Orbiter entry trajectory considerations

Space shuttle trajectory-shaping optimization considering the vehicle's thermal environment and the resulting requirements of the thermal protection system (TPS) are discussed. Optimization studies have been conducted yielding results which are generally applicable to shuttle orbiters and are independent of evolution and redirection of the space shuttle program. Two investigations of optimal trajectory shaping, with different methods of considering the thermal environment are presented.

Rehder, J. J.↗

NASA NTRS1993

Exploiting orbital effects for short-range extravehicular transfers

The problem studied in this paper is that of using Simplified Aid for Extravehicular Activity (EVA) Rescue (SAFER) to carry out efficient short-range transfers from the payload bay of the Space Shuttle Orbiter to the vicinity of the underside of the vehicle, for instance for inspection and repair of thermal tiles or umbilical doors. Trajectories are shown to exist, for the shuttle flying noise forward and belly down, that take the astronaut to the vicinity of the underside with no thrusting after the initial push-off. However, these trajectories are too slow to be of practical interest, as they take roughly an hour to execute. Additionally, they are quite sensitive to errors in the initial push-off rates. To overcome both of these difficulties, trajectories are then studied which include a single in-flight impulse of small magnitude ( in the range 0.1 - 0.4 fps). For operational simplicity, this impulse is applied towards the Orbiter at the moment when the line-of -sight of the EVA crewmember is tangential to the underside of the vehicle. These trajectories are considerably faster than the non-impulsive ones: transit times of less than 10 minutes are achievable. Furthermore, the man-in-the-loop feedback scheme used for impulse timing greatly reduces the sensitivity to initial velocity errors. Finally, similar one-impulse trajectories are also shown to exist for the Orbiter in a gravity-gradient attitiude.

Williams, Trevor↗

NASA NTRS2011

Ares I-X Range Safety Analyses Overview

Ares I-X was the first test flight of NASA's Constellation Program's Ares I Crew Launch Vehicle designed to provide manned access to low Earth orbit. As a one-time test flight, the Air Force's 45th Space Wing required a series of Range Safety analysis data products to be developed for the specified launch date and mission trajectory prior to granting flight approval on the Eastern Range. The range safety data package is required to ensure that the public, launch area, and launch complex personnel and resources are provided with an acceptable level of safety and that all aspects of prelaunch and launch operations adhere to applicable public laws. The analysis data products, defined in the Air Force Space Command Manual 91-710, Volume 2, consisted of a nominal trajectory, three sigma trajectory envelopes, stage impact footprints, acoustic intensity contours, trajectory turn angles resulting from potential vehicle malfunctions (including flight software failures), characterization of potential debris, and debris impact footprints. These data products were developed under the auspices of the Constellation's Program Launch Constellation Range Safety Panel and its Range Safety Trajectory Working Group with the intent of beginning the framework for the operational vehicle data products and providing programmatic review and oversight. A multi-center NASA team in conjunction with the 45th Space Wing, collaborated within the Trajectory Working Group forum to define the data product development processes, performed the analyses necessary to generate the data products, and performed independent verification and validation of the data products. This paper outlines the Range Safety data requirements and provides an overview of the processes established to develop both the data products and the individual analyses used to develop the data products, and it summarizes the results of the analyses required for the Ares I-X launch.

Starr, Brett R.↗

NASA NTRS2011

Ares I-X Range Safety Flight Envelope Analysis

Starr, Brett R.↗

NASA NTRSDate not supplied

Mars Entry, Descent, and Landing Instrumentation 2 Trajectory, Aerodynamics, and Atmosphere Reconstruction

On February 18th, 2021, the Mars 2020 entry system successfully delivered the Perseverance rover to the surface of Mars at Jezero Crater. The entry capsule carried a set of instrumentation installed on the heat shield and backshell, named the Mars Entry, Descent, and Landing Instrumentation 2. The instruments include pressure transducers, thermocouples, heat flux gauges, and radiometers to measure the aerodynamic and aerothermodynamic performance of the entry vehicle. This paper describes the trajectory and atmosphere reconstruction results based on the pressure sensor measurements. The process uses a Kalman filter approach to estimate the freestream atmospheric properties from the pressure measurements combined with a model of the pressure distribution of the heatshield and other sensor inputs, including an inertial measurement unit and other on-board navigation sensors, and several external atmospheric observations. The results indicate upper altitude density was up to 150% higher than nominal, which is consistent with the observed early entry guidance start time. The density below 40 km was within 12% the pre-flight predictions. The reconstructed axial force coefficient was approximately 2% lower than the pre-flight prediction across the flight range.

Christopher D Karlgaard↗

NASA NTRS2007

Reentry Thermal Analysis of a Generic Crew Exploration Vehicle Structure

Comparative studies were performed on the heat-shielding characteristics of honeycomb-core sandwich panels fabricated with different materials for possible use as wall panels for the proposed crew exploration vehicle. Graphite/epoxy sandwich panel was found to outperform aluminum sandwich panel under the same geometry due to superior heat-shielding qualities and lower material density. Also, representative reentry heat-transfer analysis was performed on the windward wall structures of a generic crew exploration vehicle. The Apollo low Earth orbit reentry trajectory was used to calculate the reentry heating rates. The generic crew exploration vehicle has a graphite/epoxy composite honeycomb sandwich exterior wall and an aluminum honeycomb sandwich interior wall, and is protected with the Apollo thermal protection system ablative material. In the thermal analysis computer program used, the TPS ablation effect was not yet included; however, the results from the nonablation heat-transfer analyses were used to develop a "virtual ablation" method to estimate the ablation heat loads and the thermal protection system recession thicknesses. Depending on the severity of the heating-rate time history, the virtual ablation period was found to last for 87 to 107 seconds and the ablation heat load was estimated to be in the range of 86 to 88 percent of the total heat load for the ablation time period. The thermal protection system recession thickness was estimated to be in the range of 0.08 to 0.11 inches. For the crew exploration vehicle zero-tilt and 18-degree-tilt stagnation points, thermal protection system thicknesses of h = {0.717, 0.733} inches were found to be adequate to keep the substructural composite sandwich temperature below the limit of 300 F.

Ko, William L.↗

NASA NTRS2021

An Approach to Reasoning Service Migration in Data and Reasoning Fabric (DRF) Implementation

In this paper we consider service migration problem for Data and Reasoning Fabric (DRF) enabled airspace operations assuming a fixed cloud/edge infrastructure with allocated computing, storage and power resources, where cloud/edge servers, and communication stations are in a wired connected network, while vehicles use a wireless network for communication. The objective is to automatically select the best location for the requested service execution, which achieves minimum cost while satisfying the user quality of service (QoS) and available resources constraints. To this end, estimates of the response time, consumed energy and total cost are defined for each potential compute location. A mixed-integer linear program is then formulated and solved to identify optimal compute locations given QoS constraints, network infrastructure limitations, with worst-case vehicle positioning. The approach is applied to trajectory re-planning use case to avoid a collision with an emergency vehicle in real time.

Air mobility↗

NASA NTRS1997

Lunar Prospector Mission Design

The National Aeronautics and Space Administration (NASA) has selected Lunar Prospector as one of the discovery missions to conduct solar system exploration science investigations. The mission is NASA's first lunar voyage to investigate key science objectives since Apollo and will be launched in September 1997. In keeping with discovery program requirements to reduce total mission cost and utilize new technology, Lunar Prospector's mission design and control will focus on the use of innovative and proven trajectory analysis programs. As part of this effort, the Ames Research Center and the Goddard Space Flight Center have become partners in the Lunar Prospector trajectory team to provide the trajectory analysis and orbit determination support. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with an injection into a circular 100 km altitude polar mapping orbit. The mission design will incorporate the optimization of the reference trajectory for a new launch vehicle. An extended mission includes a periselene altitude approaching 10 km. As part of the navigation, a new lunar potential model will be used that incorporates data from the recent Clementine mission. This paper provides understanding of the design of the Lunar Prospector mission and the optimization process of the nominal trajectory, orbit determination, launch windows, and lunar orbit targets, and discusses the feasibility of meeting the mission goals.

Folta, David↗

NASA NTRS1985

Candidate techniques for non-intrusive measurement of flight boundary-layer properties

Techniques to measure gas composition and chemical activity in the hypersonic boundary layer associated with entry research vehicles and with a tethered satellite deployed at 100-125 km altitude are currently being investigated at the NASA-Langley Research Center. These measurements are appropriate for the study of the aerothermodynamic properties of these vehicles when operating with controlled trajectories in the thermosphere between 125 and 55 km. A number of the required quantitative physical measurements have been identified, and candidate techniques to obtain the measurements without perturbing the flow are being considered. Two candidate techniques are discussed in this paper: a mass spectrometer to analyze the gas composition at the vehicle surface, and Rayleigh scattering to obtain density, velocity, and temperature gradients from the surface to the edge of the boundary layer.

Wood, G. M., Jr.↗

NASA NTRS1984

Minimum energy-loss guidance for aero-assisted orbital plane change

Minimum energy-loss guidance for the aero-assisted plane change of an orbiting vehicle is developed and applied to the plane change of a circular orbit. First, trajectories which minimize the fuel required to change the orbital plane are computed for a realistic vehicle. From these trajectories, it is observed that the fuel weight is minimized if the velocity at exit from the atmosphere is maximized. Next, for the atmospheric turn, approximate optimal controls (angle of attack and bank angle) which maximize the exit velocity are derived. Finally, the minimum-fuel problem is resolved using optimal guidance for the atmospheric part of the trajectory, and the optimization problem reduces to a one-dimensional parameter minimization. Successful plane changes up to 40 deg are demonstrated. Optimal guidance requires up to 14 percent more fuel than the 'true' optimum but only 50 percent of the fuel required by the single-impulse maneuver. Finally, the guidance law developed here is implementable because only algebraic manipulations are required.

Hull, D. G.↗

NASA NTRS1993

Profile negotiation: An air/ground automation integration concept for managing arrival traffic

NASA Ames Research Center and NASA Langley Research Center conducted a joint simulation study to evaluate a profile negotiation process (PNP) between a time-based air traffic control ATC system and an airplane equipped with a four dimensional flight management system (4D FMS). Prototype procedures were developed to support the functional implementation of this process. The PNP was designed to provide an arrival trajectory solution that satisfies the separation requirements of ATC while remaining as close as possible to the airplane's preferred trajectory. The Transport Systems Research Vehicle cockpit simulator was linked in real-time to the Center/TRACON Automation System (CTAS) for the experiment. Approximately 30 hours of simulation testing were conducted over a three week period. Active airline pilot crews and active Center controller teams participated as test subjects. Results from the experiment indicate the potential for successful incorporation of airplane preferred arrival trajectories in the CTAS automation environment. Controllers were able to consistently and effectively negotiate nominally conflict-free trajectories with pilots flying a 4D-FMS-equipped airplane. The negotiated trajectories were substantially closer to the airplane's preference than would have otherwise been possible without the PNP. Airplane fuel savings relative to baseline CTAS were achieved in the test scenarios. The datalink procedures and clearances developed for this experiment, while providing the necessary functionality, were found to be operationally unacceptable to the pilots. Additional pilot control and understanding of the proposed airplane-preferred trajectory and a simplified clearance procedure were cited as necessary for operational implementation of the concept. From the controllers' perspective, the main concerns were the ability of the 4D airplane to accurately track the negotiated trajectory and the workload required to support the PNP as implemented in this study.

Williams, David H.↗

NASA NTRS1977

Space shuttle engineering and operations support. ALT separation reference trajectories for tailcone on orbiter forward and aft CG configurations. Mission planning, mission analysis and software formulation

A preflight analysis of the ALT separation reference trajectories for the tailcone on, forward, and aft cg orbiter configurations is documented. The ALT separation reference trajectories encompass the time from physical separation of the orbiter from the carrier to orbiter attainment of the maximum ALT interface airspeed. The trajectories include post separation roll maneuvers by both vehicles and are generated using the final preflight data base. The trajectories so generated satisfy all known separation design criteria and violate no known constraints. The requirement for this analysis is given along with the specifications, assumptions, and analytical approach used to generate the separation trajectories. The results of the analytical approach are evaluated, and conclusions and recommendations are summarized.

Glenn, G. M.↗

NASA NTRS2000

A Study of a Lifting Body as a Space Station Crew Exigency Return Vehicle (CERV)

A lifting body is described for use as a return vehicle for crews from a space station. Reentry trajectories, subsystem weights and performance, and costs are included. The baseline vehicle is sized for a crew of eight. An alternate configuration is shown in which only four crew are carried with the extra volume reserved for logistics cargo. A water parachute recovery system is shown as an emergency alternative to a runway landing. Primary reaction control thrusters from the Shuttle program are used for orbital maneuvering while the Shuttle verniers are used for all attitude control maneuvers.

MacConochie, Ian O.↗

NASA NTRS1966

MSFC Flight Mission Directive Apollo-Saturn 205 Mission

The purpose of this directive is to provide, under one cover, coordinated direction for the AS-205 Space Vehicle Flight. Within this document, mission objectives are specified, vehicle configuration is described and referenced, flight trajectories, data acquisition requirements, instrumentation requirements, and detailed documentation requirements necessary to meet launch vehicle mission objectives are defined and/or referenced.

Source record↗