Results for “INTERPLANETARY TRAJECTORY” — Page 9

NASA NTRS2000

A Flight Demonstration of Plasma Rocket Propulsion

The Advanced Space Propulsion Laboratory at the NASA Johnson Space Center has been engaged in the development of a variable specific impulse magnetoplasma rocket (V ASIMR) for several years. This type of rocket could be used in the future to propel interplanetary spacecraft and has the potential to open the entire solar system to human exploration. One feature of this propulsion technology is the ability to vary its specific impulse so that it can be operated in a mode that maximizes propellant efficiency or a mode that maximizes thrust. Variation of specific impulse and thrust enhances the ability to optimize interplanetary trajectories and results in shorter trip times and lower propellant requirements than with a fixed specific impulse. In its ultimate application for interplanetary travel, the VASIMR would be a multi-megawatt device. A much lower power system is being designed for demonstration in the 2004 timeframe. This first space demonstration would employ a lO-kilowatt thruster aboard a solar powered spacecraft in Earth orbit. The 1O-kilowatt V ASIMR demonstration unit would operate for a period of several months with hydrogen or deuterium propellant with a specific impulse of 10,000 seconds.

Petro, Andrew↗

NASA NTRS2017

From Basking Ridge to the Jupiter Trojans

This presentation describes the activities of the Global Trajectory Optimization Lab, a subdivision of the Navigation and Mission Design Branch at NASA GSFC. The students will learn the basics of interplanetary trajectory optimization and then, as an example, the Lucy mission to the Jupiter Trojans will be described from both a science and engineering perspective.

trajectory design↗

NASA NTRS2020

Using Reinforcement Learning to Design Missed Thrust Resilient Trajectories

From ion thrusters to solar sails, spacecraft continue to adopt new and more efficient forms of propulsion. As these low-thrust propulsion meth- ods have become more prevalent, new challenges have arisen. Depending on the mission, low-thrust propulsion elements may need to thrust con- tinuously for days/months. During these thrusting periods, external fac- tors, such as a micro-meteoroid impact or a software glitch, may cause the spacecraft to prematurely cease its thrust stage. Half of all deep space missions enter a safe mode where they cannot thrust every four months. These missed thrust events can result in the complete loss of a space- craft for time-dependent trajectories like planetary rendezvous. This paper demonstrates how neural networks, trained using reinforcement learning, can autonomously correct for missed thrust events during an interplanetary trajectory.

Laipert, Frank E.↗

NASA NTRS2013

Interplanetary Departure Stage Navigation by Means of Liaison Orbit Determination Architecture

Autonomous orbit determination for departure stages of interplanetary trajectories is conducted by means of realistic radiometric observations between the departing spacecraft and a satellite orbiting the first lunar libration point. Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) is used to estimate the orbit solution. This paper uses high-fidelity simulations to explore the utilization of LiAISON in providing improved accuracy for interplanetary departure missions. The use of autonomous navigation to supplement current techniques for interplanetary spacecraft is assessed using comparisons with groundbased navigation. Results from simulations including the Mars Science Laboratory, Mars Exploration Rover, and Cassini are presented. It is shown that observations from a dedicated LiAISON navigation satellite could be used to supplement ground-based measurements and significantly improve tracking performance.

interplanetary↗

NASA NTRS2013

Interplanetary Departure Stage Navigation by Means of Liaison Orbit Determination Architecture

Autonomous orbit determination for departure stages of interplanetary trajectories is conducted by means of realistic radiometric observations between the departing spacecraft and a satellite orbiting the first lunar libration point. Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) is used to estimate the orbit solution. This paper uses high-fidelity simulations to explore the utilization of LiAISON in providing improved accuracy for interplanetary departure missions. The use of autonomous navigation to supplement current techniques for interplanetary spacecraft is assessed using comparisons with groundbased navigation. Results from simulations including the Mars Science Laboratory, Mars Exploration Rover, and Cassini are presented. It is shown that observations from a dedicated LiAISON navigation satellite could be used to supplement ground-based measurements and significantly improve tracking performance.

Linked Autonomous Interplanetary Satellite Orbit N↗

NASA NTRS2009

Broad Search Solar Electric Propulsion Trajectories to Saturn with Gravity Assists

Solar electric propulsion (SEP) trajectories to Saturn using multiple gravity assists are explored for the joint NASA and ESA Titan Saturn System Mission study. Results show that these new trajectories enable greater performance compared to chemical propulsion with similar gravity assists or SEP without gravity assists. This paper describes the method used in finding these interplanetary trajectories and examines variations in the performance for different SEP systems, flight times, and flyby sequences. The benefits of the SEP trajectories for a mission to Saturn are also discussed.

Titan↗

NASA NTRS2001

Development of a Mars Airplane Entry, Descent, and Flight Trajectory

An entry, descent, and flight (EDF) trajectory profile for a Mars airplane mission is defined as consisting of the following elements: ballistic entry of an aeroshell; supersonic deployment of a decelerator parachute; subsonic release of a heat shield; release, unfolding, and orientation of an airplane to flight attitude; and execution of a pull up maneuver to achieve trimmed, horizontal flight. Using the Program to Optimize Simulated Trajectories (POST) a trajectory optimization problem was formulated. Model data representative of a specific Mars airplane configuration, current models of the Mars surface topography and atmosphere, and current estimates of the interplanetary trajectory, were incorporated into the analysis. The goal is to develop an EDF trajectory to maximize the surface-relative altitude of the airplane at the end of a pull up maneuver, while subject to the mission design constraints. The trajectory performance was evaluated for three potential mission sites and was found to be site-sensitive. The trajectory performance, examined for sensitivity to a number of design and constraint variables, was found to be most sensitive to airplane mass, aerodynamic performance characteristics, and the pull up Mach constraint. Based on the results of this sensitivity study, an airplane-drag optimized trajectory was developed that showed a significant performance improvement.

Murray, James E.↗

NASA NTRS1992

Galileo trajectory design

The trajectory design of the Galileo spacecraft is examined. The Galileo spacecraft was launched on a six-year long trip to Jupiter in October 1989. A new Venus-Earth-Earth-Gravity Assist (VEEGA) trajectory mode is being used for the transfer to Jupiter and involves two phasing orbits around the sun and gravity-assist flybys with Venus. The aggregate delta V acquired from these flybys is 18.3 km/s. The interplanetary trajectory includes a close flyby of asteroid 951-Gaspra in October 1991 and a possible flyby of 243-Ida in August 1993. After arrival at Jupiter in December 1995, the previously released Galileo atmospheric probe will relay data to earth via the Galileo Orbiter. The orbital phase of the mission will involve 10 orbits of Jupiter over a 22 month period. In this phase the Orbiter will use repeated gravity-assisted flybys of Europa, Ganymede and Callisto during which Jupiter, its magnetosphere and the Galilean satellites will be investigated. The mission is scheduled to end in October 1997.

D'Amario, Louis A.↗

NASA NTRS1970

Parametric design of interplanetary and orbital trajectories with examples for Mars 1973, 1975, and 1977 opportunities

Hampshire, W. F., II↗

NASA NTRS2012

Time and Energy, Exploring Trajectory Options Between Nodes in Earth-Moon Space

The Global Exploration Roadmap (GER) was released by the International Space Exploration Coordination Group (ISECG) in September of 2011. It describes mission scenarios that begin with the International Space Station and utilize it to demonstrate necessary technologies and capabilities prior to deployment of systems into Earth-Moon space. Deployment of these systems is an intermediate step in preparation for more complex deep space missions to near-Earth asteroids and eventually Mars. In one of the scenarios described in the GER, "Asteroid Next", there are activities that occur in Earth-Moon space at one of the Earth-Moon Lagrange (libration) points. In this regard, the authors examine the possible role of an intermediate staging point in an effort to illuminate potential trajectory options for conducting missions in Earth-Moon space of increasing duration, ultimately leading to deep space missions. This paper will describe several options for transits between Low Earth Orbit (LEO) and the libration points, transits between libration points, and transits between the libration points and interplanetary trajectories. The solution space provided will be constrained by selected orbital mechanics design techniques and physical characteristics of hardware to be used in both crewed missions and uncrewed missions. The relationships between time and energy required to transfer hardware between these locations will provide a better understanding of the potential trade-offs mission planners could consider in the development of capabilities, individual missions, and mission series in the context of the ISECG GER.

Martinez, Roland↗

NASA NTRS1991

SeGRAm - A practical and versatile tool for spacecraft trajectory optimization

An implementation of the Sequential Gradient/Restoration Algorithm, SeGRAm, is presented along with selected examples. This spacecraft trajectory optimization and simulation program uses variational calculus to solve problems of spacecraft flying under the influence of one or more gravitational bodies. It produces a series of feasible solutions to problems involving a wide range of vehicles, environments and optimization functions, until an optimal solution is found. The examples included highlight the various capabilities of the program and emphasize in particular its versatility over a wide spectrum of applications from ascent to interplanetary trajectories.

Rishikof, Brian H.↗

NASA NTRS2017

NASA Double Asteroid Redirection Test (Dart) Trajectory Validation and Robustness

The Double Asteroid Redirection Test (DART) mission will be the first to test the concept of a kinetic impactor. Several studies have been made on asteroid redirection and impact mitigation, however, to this date no mission tested the proposed concepts. An impact study on a representative body allows the measurement of the effects on the target's orbit and physical structure. With this goal, DART's objective is to verify the effectiveness of the kinetic impact concept for planetary defense. The spacecraft uses solar electric propulsion to escape Earth, flyby (138971) 2001 CB21 for impart rehearsal, and impact the secondary body of the (65803) Didymos system. This work focuses on the interplanetary trajectory design part of the mission with the validation of the baseline trajectory, performance comparison to other mission objectives, and assessment of the baseline robustness to missed thrust events. Results show a good performance of the selected trajectory for different mission objectives: latest possible escape date, maximum kinetic energy on impact, shortest possible time of flight, and use of an Earth swing-by. The baseline trajectory was shown to be robust to a missed thrust with 1% of fuel margin being enough to recover the mission for failures of more than 14 days.

design↗

NASA NTRS1976

Preentry communication design elements for outer planets atmospheric entry probe

Four related tasks are discussed for data transmission from a probe prior to entering the atmosphere of Jupiter to an orbiting spacecraft in a trajectory past the planet: (1) link analysis and design; (2) system conceptual design; (3) Doppler measurement analysis; and (4) an electronically despun antenna. For tasks 1, 3, and 4, an analytical approach was developed and combined with computational capability available to produce quantitative results corresponding to requirements and constraints given by NASA, ARC. One constraint having a major impact on the numerical results of the link analysis was the assumption of a nonsteerable antenna on a spinning orbiter. Other constraints included the interplanetary trajectory and the approach trajectory. Because the Jupiter Orbiter Probe (JOP) program is currently in a state of evolution, all requirements and constraints applied during this study are subject to change. However, the relationships of parameters as developed will remain valid and will aid in planning Jupiter missions.

Source record↗

NASA NTRS2008

Entry, Descent, and Landing Operations Analysis for the Stardust Entry Capsule

On the morning of January 15, 2006, the Stardust capsule successfully landed at the Utah Test and Training range in northwest Utah returning cometary samples from the comet Wild-2. An overview of the entry, descent, and landing (EDL) trajectory analysis that was performed for targeting during the mission operations phase upon final approach to Earth is described. The final orbit determination solution produced an inertial entry flight-path angle of -8.21 deg (the desired nominal value) with a 3-sigma uncertainty of +/-0.0017 deg (2% of the requirement). The navigation and EDL operations effort accurately delivered the entry capsule to the desired landing site. The final landing location was 8.1 km from the target, which was well within the allowable landing area. Overall, the Earth approach operation procedures worked well and there were no issues (logistically or performance based) that arose. As a result, the process of targeting a capsule from an interplanetary trajectory and accurately landing it on Earth was successfully demonstrated.

Desai, Prasun N.↗

NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗

NASA NTRSDate not supplied

Improvements to Thermal Protection System Design of Aerocapture Systems for Uranus Orbiters

The National Academies Planetary Science and Astrobiology Decadal Survey identified Uranus and Neptune - called Ice Giants - as the priority destinations for science. The survey assessed both a mission to Uranus through the Uranus Orbiter and Probe (UOP) concept, and Neptune through the Neptune-Triton Odyssey concept and determined that Uranus is the highest priority for a Flagship class mission. The UOP mission concept planned to deliver an in situ probe and conduct a multi-year orbital tour of the system to meet the science objectives. While the Uranus mission is currently viable with launch windows starting in 2031 using existing launch vehicles, the mission has a cruise phase of at least 12 years and would require more than half of its weight in fuel propellant to achieve the change in velocity necessary for orbital insertion. Aerocapture uses aerodynamic forces generated on a vehicle by the planet's atmosphere to modulate a spacecraft's trajectory, decreasing spacecraft velocity, and allowing mission designers to target the final orbital state. Aerocapture reduces the time-of-flight from Earth to Uranus over a fully propulsive solution, opening up more launch opportunities to arrive in the 2040's to the mission's science objectives. Aerocapture also allows a payload mass increase by mitigating the need for fuel to retropropulsively insert the payload into orbit, thereby increase the science that can be performed. For an aerocapture mission structure using a traditional aeroshell to deliver the UOP scientific payload to Uranus, Conformal Phenolic Impregnated Carbon Ablator (C-PICA) was determined to be the best-performing forebody thermal protection system (TPS) and other candidate aftbody TPS options were presented as feasible. This paper focuses on A) evaluating C-PICA as a forebody TPS using stressing entry conditions associated with a large range of potential Uranus flagship launch vehicles and interplanetary trajectories, B) widening the aftbody TPS candidates for new mass-efficient and cost-efficient solutions, and C) a summary list of actions remaining to provide a technically feasible and supply-robust set of TPS for an aerocapture vehicle to the Ice Giants is presented.

C-PICA↗

NASA NTRS2020

Aerocapture as an Option for Ice Giants Mission

Aerocapture is an atmospheric maneuver where the aerodynamic forces of the vehicle (lift and drag) are used to provide the 𝞓V needed to slow down from the approach hyperbolic trajectory to achieve the desired captured orbit around the target planet. The aeroassist capture provides a large savings in propulsion needed to change the velocity of the vehicle, since aerodynamic forces rather than propulsive systems provide the change in velocity. Aerocapture requires an integrated system level design, including thermal protection systems, actuator systems for aerodynamic modulation, and guidance and control systems that can autonomously command the change in the aeroassist forces. Although aerocapture has been proposed for many situations in the past, including Mars, Venus, Titan, Saturn, Uranus, and Neptune orbiters as well as for Earth demonstration missions, it has not been attempted on any missions. However, many studies in the past three decades that have considered aerocapture as a design option have concluded that there are large mass savings that come from using aeroassist forces rather than propulsive forces to put a spacecraft into a captured orbit. The benefits are destination dependent, but some of the largest mass savings occur for the Ice Giants planets. Due to the large hyperbolic velocities of interplanetary trajectories approaching Uranus and Neptune, large amount of propulsion must be used to put a spacecraft in science orbits around these planets. Aerocapture can reduce the propulsion needs by dissipating energy in the sizable atmospheres of Uranus and Neptune. NASA commissioned a detailed study analysis to quantify the benefits, if any, for use at Neptune. The study found that even having to provide a heat shield aerocapture could deliver 40% more payload than an all-propulsive vehicle, and also provides for a 3–4-year reduction in trip time. Mass savings are expected also at Uranus with an aerocapture mission. Additionally, more recent advances in thermal protection systems and guidance and control systems show a path to increase capabilities beyond those results. This paper will discuss the merits of including aerocapture as an option for an Ice Giants mission. The discussion will focus on the cost and mass savings of using aerocapture instead of propulsive burns for Ice Giants orbiter mission scenarios while also detailing a potential concept of operations and entry vehicle design. Finally, the talk will discuss more recent work that shows aerocapture design is possible without the development of any novel entry vehicle. Thus, aerocapture can significantly improve science capabilities for an Ice Giants mission.

Soumyo Dutta↗

NASA NTRS2008

Mission and Navigation Design for the 2009 Mars Science Laboratory Mission

NASA s Mars Science Laboratory mission will launch the next mobile science laboratory to Mars in the fall of 2009 with arrival at Mars occurring in the summer of 2010. A heat shield, parachute, and rocket-powered descent stage, including a sky crane, will be used to land the rover safely on the surface of Mars. The direction of the atmospheric entry vehicle lift vector will be controlled by a hypersonic entry guidance algorithm to compensate for entry trajectory errors and counteract atmospheric and aerodynamic dispersions. The key challenges for mission design are (1) develop a launch/arrival strategy that provides communications coverage during the Entry, Descent, and Landing phase either from an X-band direct-to-Earth link or from a Ultra High Frequency link to the Mars Reconnaissance Orbiter for landing latitudes between 30 deg North and 30 deg South, while satisfying mission constraints on Earth departure energy and Mars atmospheric entry speed, and (2) generate Earth-departure targets for the Atlas V-541 launch vehicle for the specified launch/arrival strategy. The launch/arrival strategy employs a 30-day baseline launch period and a 27-day extended launch period with varying arrival dates at Mars. The key challenges for navigation design are (1) deliver the spacecraft to the atmospheric entry interface point (Mars radius of 3522.2 km) with an inertial entry flight path angle error of +/- 0.20 deg (3 sigma), (2) provide knowledge of the entry state vector accurate to +/- 2.8 km (3 sigma) in position and +/- 2.0 m/s (3 sigma) in velocity for initializing the entry guidance algorithm, and (3) ensure a 99% probability of successful delivery at Mars with respect to available cruise stage propellant. Orbit determination is accomplished via ground processing of multiple complimentary radiometric data types: Doppler, range, and Delta-Differential One-way Ranging (a Very Long Baseline Interferometry measurement). The navigation strategy makes use of up to five interplanetary trajectory correction maneuvers to achieve entry targeting requirements. The requirements for cruise propellant usage and atmospheric entry targeting and knowledge are met with ample margins.

Mission design↗