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At least 271 records · Page 15

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NASA NTRS2015

Lunar Cube Transfer Trajectory Options

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can be considered which have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfers and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

Libration Orbits↗

NASA NTRS2015

Lunar Cube Transfer Trajectory Options

Numerous Earth-Moon trajectory and lunar orbit options are available for Cubesat missions. Given the limited Cubesat injection infrastructure, transfer trajectories are contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these transfers can be restricted by the selection or designs of Cubesat subsystems such as propulsion or communication. Nonetheless, many trajectory options can b e considered which have a wide range of transfer duration, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several options including deployment from low Earth orbit (LEO) geostationary transfer orbits (GTO) and higher energy direct lunar transfer and the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory.

Libration Orbits↗

NASA NTRS2015

OSIRIS-REx, Returning the Asteroid Sample

This paper addresses the technical aspects of the sample return system for the upcoming Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission. The overall mission design and current implementation are presented as an overview to establish a context for the technical description of the reentry and landing segment of the mission.The prime objective of the OSIRIS-REx mission is to sample a primitive, carbonaceous asteroid and to return that sample to Earth in pristine condition for detailed laboratory analysis. Targeting the near-Earth asteroid Bennu, the mission launches in September 2016 with an Earth reentry date of September 24, 2023.OSIRIS-REx will thoroughly characterize asteroid Bennu providing knowledge of the nature of near-Earth asteroids that is fundamental to understanding planet formation and the origin of life. The return to Earth of pristine samples with known geologic context will enable precise analyses that cannot be duplicated by spacecraft-based instruments, revolutionizing our understanding of the early Solar System. Bennu is both the most accessible carbonaceous asteroid and one of the most potentially Earth-hazardous asteroids known. Study of Bennu addresses multiple NASA objectives to understand the origin of the Solar System and the origin of life and will provide a greater understanding of both the hazards and resources in near-Earth space, serving as a precursor to future human missions to asteroids.This paper focuses on the technical aspects of the Sample Return Capsule (SRC) design and concept of operations, including trajectory design and reentry retrieval. Highlights of the mission are included below.The OSIRIS-REx spacecraft provides the essential functions for an asteroid characterization and sample return mission: attitude control propulsion power thermal control telecommunications command and data handling structural support to ensure successful rendezvous with Bennu characterization of Bennus properties delivery of the sampler to the surface, and return of the spacecraft to the vicinity of the Earth sample collection, performed by the Touch-and-Go Sample Acquisition Mechanism (TAGSAM), to acquire a regolith sample from the surface Earth re-entry and SRC recovery. Following sample collection, OSIRIS-REx drifts away from Bennu until the Asteroid Departure Maneuver is commanded on March 4, 2021, sending OSIRIS-REx on a ballistic return cruise to Earth. No additional large deterministic maneuvers are required to return the SRC to Earth. During the cruise, tracking and trajectory correction maneuvers (TCMs) are performed as necessary to precisely target the entry corridor. As OSIRIS-REx approaches Earth, the reentry plans are reviewed starting about a year before arrival, and preparations begin. The spacecraft is targeted away from the Earth until 7 days before entry. The final two trajectory correction maneuvers bring the spacecraft on target toward the Utah Test and Training Range (UTTR), with sufficient time for contingency resolution. The SRC releases 4 hours prior to atmospheric entry interface and, using the Stardust capsule heritage design, employs a traditional drogue and main parachute descent system for a soft touchdown.

reentry↗

NASA NTRS2019

A Titan gravity-assist technique for ballistic tours skimming over the rings of Saturn

A novel type of Titan-flyby orbits featuring passes of the Saturn rings at close range has been recently discovered. The purpose of this study is to explore the trajectory design space and assess the applicability of such orbits to the design of a Saturn Ring Tour mission concept. A set of initial conditions required to start a tour of the rings is first determined and a flyby sequence to maximize the science observation time over desired ring regions in terms of range distance, relative velocity and duration is then selected. To demonstrate the potential of this technique, a sample high-fidelity ballistic ring tour is detailed.

Tiscareno, Matthew↗

NASA NTRS1976

Launch opportunity classification of VEGA and delta V-EGA trajectories to the outer planets

Two concepts in spacecraft trajectory design which have been proposed by Hollenbeck (1975) and Bender and Friedlander (1975) are considered. Both concepts are intended for planetary missions and require an earth reencounter created by either a Venus gravity assist (VEGA) or a large midcourse impulse (delta V-EGA). In the case of outer planet missions, the new concepts offer significantly lower launch energy requirements than the direct ballistic flight mode. However, at least an additional two years flight time is needed for geometry phasing to set up the earth reencounter. A description is given of the trajectory characteristics for the two concepts. Attention is also given to the method of analysis used for the studies, aspects of trajectory generation, payload calculations, and target opportunity data.

Stancati, M. L.↗

NASA NTRS1974

Mission design for a ballistic slow flyby Comet Encke 1980

Preliminary mission analyses for a proposed 1980 slow flyby (7-9 km/s) of comet Encke are presented. Among the topics covered are science objectives, Encke's physical activity and ephemeris accuracy, trajectory and launch-window analysis, terminal guidance, and spacecraft concepts. The nominal mission plan calls for a near-perihelion intercept with two spacecraft launched on a single launch vehicle. Both spacecraft will arrive at the same time, one passing within 500 km from Encke's nucleus on its sunward side, the other cutting through the tail region. By applying a small propulsive correction about three weeks after the encounter, it is possible to retarget both spacecraft for a second Encke intercept in 1984. The potential science return from the ballistic slow flyby is compared with other proposed mission modes for the 1980 Encke flyby mission, including the widely advocated slow flyby using solar-electric propulsion. It is shown that the ballistic slow flyby is superior in every respect.

Farquhar, R. W.↗

NASA NTRS1958

Hypersonic Flight and the Re-Entry Problem: The Twenty-First Wright Brothers Lecture

Up to this point the discussion of the problems of rocket vehicles has been confined to the effects of phenomena which have in the past been important ones for lower speed aircraft and will continue to be important for aircraft of all speeds. Now with considerable extension of both speed and altitude, other phenomena also become important. The nature of some problems will be altered, as a result, and new problems, of course, will be encountered. First, it is well to note that our interest in bluff bodies for ballistic vehicles in particular, and in rounded-nosed bodies generally, has changed our emphasis in aerodynamics. The detached bow waves which occur with such bodies at high supersonic speeds complicate the calculations of the flow-field characteristics. In the present period, much attention is being given to such studies. In addition, at the very high altitudes attained by most of the rocket-craft, the mean free path of air molecules can be of the same order, or long, compared to the dimensions of the vehicles. Thus, slip-flow and free-molecule-flow studies are of interest, particularly for satellite vehicles." The aerodynatnicist must deal with air having unfamiliar states and properties. Second, at hypersonic speeds where, for example, air is greatly decelerated, it may undergo considerable change in composition, the degree of change depending upon many factors. Dissociation of oxygen and nitrogen molecules can occur and, in addition, thermal ionization of many of the constituents. It is naturally to be expected that the convective heat transfer will, as a result, be altered from what it was for the "perfect" gas, and this has been the subject of much recent research effort. Moreover, the decelerated gas becomes capable of radiating energy and the radiative heat transfer must generally be considered for hypersonic vehicles, particularly for long-range ballistic rockets. It is not only the aerodynamic heating problems that are affected. The fact that at very high air temperature the gas becomes electrically conductive introduces new problems in radio wave transmission and reception. In addition, a conducting gas flow can, of course, be influenced by a magnetic field. The study of such flows, which has been termed "magneto gas dynamics," is still in too primitive a state to indicate how important a role it can play, but many interesting possibilities suggest themselves. Third, our experience with airplanes powered by air-breathing engines has naturally been restricted to the stratosphere, or lower. Our ignorance increases with altitude. For rockets, literally, "the sky's the limit," and it is not surprising that a great emphasis has now been placed on obtaining a more thorough understanding of the whole atmosphere. These studies are not aimed at an understanding of the chemical and physical characteristics alone, but also of the occurrence of high-energy particles, from meteors to cosmic rays, and the nature of the problems they will promote.

Wright Brother↗

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 NTRS1974

Mission design for a ballistic slow-flyby of comet Encke 1980

Preliminary mission analyses for a proposed 1980 slow flyby (7 to 9 km/sec) of comet Encke are presented. Among the topics covered are: science objectives, Encke's physical activity and ephemeris accuracy, trajectory and launch-window analysis, terminal guidance, and spacecraft concepts. The nominal mission plan calls for a near-perihelion intercept with two spacecraft using a single launch vehicle. Both spacecraft arrive at the same time, one passing within 500 km from Encke's nucleus on its sunward side, the other cutting through the tail region. By applying a small propulsive correction about three weeks after the encounter, it is possible to retarget both spacecraft for a second Encke intercept in 1984. The potential science return from the ballistic slow flyby is compared with other proposed mission modes for the 1980 Encke flyby mission including the widely advocated slow flyby using solar-electric propulsion. It is shown that the ballistic slow flyby is superior in every respect.

Farquhar, R. W.↗

NASA NTRS2017

Free-Flight Trajectory Simulation of the ADEPT Sounding Rocket Test Using CFD

A computational study of the Adaptive Deployable Entry and Placement Technology (ADEPT) Sounding Rocket (SR-1) Test is presented using the US3D flow solver. ADEPT SR-1 is intended, in part, to assess the dynamic stability of this entry vehicle architecture. Given that no dynamic stability data exists for the ADEPT geometry, a limited ballistic range campaign has been performed to characterize the vehicle's stability characteristics pre-flight for Mach numbers between 1.21 and 2.5. Here, this data is used to assess the accuracy of US3D's free-flight CFD capability. Computed trajectories from US3D and experimental data show that the flow solver compares well in vehicle oscillation frequency, downrange distance, and oscillatory amplitude during high Mach number flight (Mavg = 2.36). For Mach numbers below 1.5, the solver under predicts total angle-of-attack by an average of 16%, but compares well in oscillatory frequency and downrange distance. Additionally, a capability for simulating the trajectory of the flight article through the atmosphere using CFD is presented. This capability couples US3D's free-flight capability to an atmosphere model that accounts for changes in free-stream density and temperature as the vehicle descends. Two simulations for the purpose of demonstrating the capability and viability of this approach are applied to SR-1 flight article, and some unique challenges are discussed.

CFD↗

NASA NTRS2013

Automating Initial Guess Generation for High Fidelity Trajectory Optimization Tools

Many academic studies in spaceflight dynamics rely on simplified dynamical models, such as restricted three-body models or averaged forms of the equations of motion of an orbiter. In practice, the end result of these preliminary orbit studies needs to be transformed into more realistic models, in particular to generate good initial guesses for high-fidelity trajectory optimization tools like Mystic. This paper reviews and extends some of the approaches used in the literature to perform such a task, and explores the inherent trade-offs of such a transformation with a view toward automating it for the case of ballistic arcs. Sample test cases in the libration point regimes and small body orbiter transfers are presented.

design process↗

NASA NTRS2019

Free-Flight CFD Simulations of Ballistic Range and Flight Tests

Free-Flight CFD capability has been implemented into the finite-volume solver US3D under the Entry Systems Modeling project. Several simulations of ballistic range experiments have been performed in order to validate the simulation software and methodology. Extension of the software to flight scale trajectories with varying freestream conditions has been carried out. Results show promising ability to predict vehicle behavior as compared to flight. Finally, a multi-body free-flight capability has been developed to generalize the single-body free-flight solver to study multiple bodies in proximal flight.

Brock, Joseph M.↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗

NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

MMOD↗

NASA NTRS2005

Aerocapture Performance Analysis of A Venus Exploration Mission

A performance analysis of a Discovery Class Venus Exploration Mission in which aerocapture is used to capture a spacecraft into a 300km polar orbit for a two year science mission has been conducted to quantify its performance. A preliminary performance assessment determined that a high heritage 70 sphere-cone rigid aeroshell with a 0.25 lift to drag ratio has adequate control authority to provide an entry flight path angle corridor large enough for the mission s aerocapture maneuver. A 114 kilograms per square meter ballistic coefficient reference vehicle was developed from the science requirements and the preliminary assessment s heating indicators and deceleration loads. Performance analyses were conducted for the reference vehicle and for sensitivity studies on vehicle ballistic coefficient and maximum bank rate. The performance analyses used a high fidelity flight simulation within a Monte Carlo executive to define the aerocapture heating environment and deceleration loads and to determine mission success statistics. The simulation utilized the Program to Optimize Simulated Trajectories (POST) that was modified to include Venus specific atmospheric and planet models, aerodynamic characteristics, and interplanetary trajectory models. In addition to Venus specific models, an autonomous guidance system, HYPAS, and a pseudo flight controller were incorporated in the simulation. The Monte Carlo analyses incorporated a reference set of approach trajectory delivery errors, aerodynamic uncertainties, and atmospheric density variations. The reference performance analysis determined the reference vehicle achieves 100% successful capture and has a 99.87% probability of attaining the science orbit with a 90 meters per second delta V budget for post aerocapture orbital adjustments. A ballistic coefficient trade study conducted with reference uncertainties determined that the 0.25 L/D vehicle can achieve 100% successful capture with a ballistic coefficient of 228 kilograms per square meter and that the increased ballistic coefficient increases post aerocapture V budget to 134 meters per second for a 99.87% probability of attaining the science orbit. A trade study on vehicle bank rate determined that the 0.25 L/D vehicle can achieve 100% successful capture when the maximum bank rate is decreased from 30 deg/s to 20 deg/s. The decreased bank rate increases post aerocapture delta V budget to 102 meters per second for a 99.87% probability of attaining the science orbit.

Starr, Brett R.↗

NASA NTRS2000

Evaluation of Mars Entry Reconstructured Trajectories Based on Hypothetical 'Quick-Look' Entry Navigation Data

A first order simulation analysis of the navigation accuracy expected from various Navigation Quick-Look data sets is performed. Here quick-look navigation data are observations obtained by hypothetical telemetried data transmitted on the fly during a Mars probe's atmospheric entry. In this simulation study, navigation data consists of 3-axis accelerometer sensor and attitude information data. Three entry vehicle guidance types are studied: I. a Maneuvering entry vehicle (as with Mars 01 guidance where angle of attack and bank angle are controlled); II. Zero angle-of-attack controlled entry vehicle (as with Mars 98); and III. Ballistic, or spin stabilized entry vehicle (as with Mars Pathfinder);. For each type, sensitivity to progressively under sampled navigation data and inclusion of sensor errors are characterized. Attempts to mitigate the reconstructed trajectory errors, including smoothing, interpolation and changing integrator characteristics are also studied.

Pastor, P. Rick↗

NASA NTRS1975

Outer planet entry nonequilibrium heating

An analytical tool has been developed which enables the impact of nonequilibrium effects on outer planet entry heating to be estimated. The analysis combines recent shock tube experiments, flow field calculations, and planetary entry trajectory analysis. The thickness of the nonequilibrium layer and its variation around the entry body have been correlated by a reaction flow parameter over a wide range of entry conditions. The influence of nonequilibrium effects on heating during entry into Saturn and Jupiter model atmospheres has been studied and the effect of vehicle size and ballistic coefficient determined. A nonequilibrium layer of significant thickness was found to exist during portions of entry into Saturn and Jupiter warm atmospheres. However, the nonequilibrium layer was found to be thin during the peak heating portions of the trajectory and resulted in reductions in total probe heating of less than 15 per cent.

Leibowitz, L. P.↗

NASA NTRSDate not supplied

Baseline Orbit Generation for Near Rectilinear Halo Orbits

A Near Rectilinear Halo Orbit (NRHO) with an eclipse-favorable phase serves as the current baseline orbit for the Gateway mission. The baseline provides a catalog of orbit maintenance targets and serves as a reference path for other types of trajectory analysis. In this investigation, an automated multiple-shooting differential corrections strategy is developed to generate many-revolution, long-duration, ballistic baseline orbits in a higher-fidelity ephemeris force model at the eclipse-avoidant phases and at other alternative phases. Additional geometric considerations for eclipse avoidance in the NRHO are presented and utilized within a long-lead orbit maintenance-based phase shift strategy.

Emily M. Zimovan-Spreen↗