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Double Asteroid Redirection Test will be the first mission to demonstrate and characterize the concept of a kinetic impactor for planetary defense, by impacting the smaller member of a binary aster...

The Mars Exploration Rover mission delivered the rovers Spirit and Opportunity to the surface of Mars using the same entry, descent, and landing scenario that was developed and successfully implemented by Mars Pathfinder. This investigation describes the premission trajectory analysis that was performed for the hypersonic portion of the Mars Exploration Rover entry up to parachute deployment. In this analysis, a six-degree-of-freedom trajectory simulation of the entry is performed to determine the entry characteristics of the capsules. In addition, a Monte Carlo dispersion analysis is also performed to statistically assess the robustness of the entry design to off-nominal conditions to ensure that all entry requirements are satisfied. The premission results show that the attitude at peak heating and parachute deployment are well within entry limits. In addition, the parachute deployment dynamic pressure and Mach number are also well within the design requirements.

Mars ascent trajectories are calculated for small-scale vehicles that would improve the affordability of Mars sample return. Vehicle size, thrust levels, staging, and the importance of atmospheric drag are all taken into consideration. The high acceleration of conventional solid rockets requires a steep trajectory for drag avoidance, followed by a relatively large circularization burn, appropriate for a second stage. Lower thrust reduces total ∆v because reduced drag permits less steep trajectories that require small circularization burns. The results suggest the development of miniature liquid-propelled vehicles or advanced solid rockets having reduced thrust and multiple-burn capability.

A new space exploration technique, called tether sling shot assist, is examined for near-Earth asteroid sample return missions. A spacecraft eying near an asteroid can attach itself to the asteroid using a space tether and an anchor device. After attachment, thevelocity vector ofthespacecraft relative totheasteroid canberotated, thereby modifying its heliocentric velocity. The link to the anchoring device can then be severed, and the spacecraft can reel the tether back for reuse. A simplieed analysis is presented of the problem, and the technical issues related to this maneuver are discussed. A trajectory analysis software is also developed and implemented to determine the possible trajectories of near-Earth asteroid sampling missions. For each trajectory, the strengths and weaknesses of different mission scenarios are highlighted. When combined with chemical propulsion, tether sling shot assist leads to very signiecant payload mass gains over conventional all-chemical propulsion systems.

The Mars Phoenix lander was launched August 4, 2007 and remained in cruise for ten months before landing in the northern plains of Mars in May 2008. The one-month Entry, Descent, and Landing (EDL) operations phase prior to entry consisted of daily analyses, meetings, and decisions necessary to determine if trajectory correction maneuvers and environmental parameter updates to the spacecraft were required. An overview of the Phoenix EDL trajectory simulation and analysis that was performed during the EDL approach and operations phase is described in detail. The evolution of the Monte Carlo statistics and footprint ellipse during the final approach phase is also provided. The EDL operations effort accurately delivered the Phoenix lander to the desired landing region on May 25, 2008.

plication to modern engineering problems/' Douglas Aircraft Co., Santa Monica, Calif. (1957); unpublished lecture notes. 8 Miele, A., "General variational theory of the flight paths of rocket powered aircraft, missiles and satellite carriers," Astronaut. Acta 4, 264-288 (1958). 9 Bauer, A. B. and Thelander, J. A., "A variational method for finding the flight path which gives the maximum burnout velocity for a multistaged rocket vehicle," Douglas Aircraft Co., Long Beach, Calif., Engineering Rept. LB-25956 (June 1959). 10 Stewart, C. E. and Thelander, J. A., "Optimum trajectory analysis," Ford Motor Co., Aeronutronic Div., Tech. Rept. 11-1435(1961).

Algorithm for trajectory analysis in closed loop terminal guidance of solar electrically thrusted interplanetary spacecraft

Mariner IV trajectory analysis and importance of flight-path selection in construction of basic mission profile

Dynamic instability of rolling high drag planetary entry vehicles at subsonic speeds, noting angle of attack divergence in computerized entry simulation

Science subsystems for Jupiter flyby missions, discussing equipment selection, mission planning, spacecraft design and trajectory and constraints on vehicle configuration

Exploration capabilities provided by Jupiter gravity assisted trajectories compared to direct ballistic flight trajectories

Mission analysis for 1972 Venus launch opportunity, discussing launch period selection and trajectory constraints for Atlas/Centaur vehicle

Spacecraft guidance analysis of multiple outer planet mission utilizing gravity assist swingbys to achieve planetary flybys with single spacecraft

Mariner Mars 1969 flyby missions objectives, experiments, spacecraft configuration, trajectory design, sterilization, flight path control and photographic imaging

Three types of Mercury missions are considered: unmanned flybys, unmanned orbiters, and manned stopovers. Velocity requirements are assessed from consideration of four transfer modes: direct, unpowered Venus swingby, modified pericenter Venus swingby, and powered Venus swingby. The total and incremental impulsive velocity requirements, the mission durations, and Earth entry velocities, where applicable, are summarized. The lowest energy trip found for each of the missions employs a form of the Venus swingby mode. The unmanned flyby mission has a minimum impulsive velocity increment (AF) of 4.2 km/sec, and the unmanned orbiter mission a minimum total AF of 11.2 km/sec. For a 1000-kg payload at Mercury, these AF's correspond to 5000 and 47,000 kg, respectively, in Earth orbit, assuming a high-pressure chemical engine. Using a low-thrust nuclear engine would reduce the weight in Earth orbit by about 10% for the orbiter mission. The manned Mercury stopovers have a minimum AFr requirement of 18.2 km/sec with a corresponding stay time of 174 days and mission duration of 414 days. These missions require an Earth orbital weight of 0.9 to 1.8 X 10 kg, depending upon the trajectory mode/nuclear engine system selected.

Round trip Mars missions combining a standard trajectory leg and a looping trajectory leg are suggested as an alternative to the more conventional trajectories. Looping trajectories are shown to make it possible to trade a slight increase in total mission time for a significant reduction in Mars stay time, with the possible added benefit of a decrease in energy requirements.

Graphical method for determining sounding rocket attitudes from optical sensor and flux gate magnetometer data

Theme T paper outlines the methods and experience obtained in applying Pattern Recognition techniques to the validation, processing, and analysis of test data in an operational environment. The first practical problem described involves trajectory analysis, prediction, and validation of models using preflight and postflight data. Two other problems relate to comparative data analysis and techniques to generate hypotheses in multidimensional problems involving manufacturing processes whereby items can be separated into success and failure groups. While the mathematical techniques presented here are standard, their use to obtain solutions to actual problems should be of interest.

Planetary trajectory handbooks for mission and system analysis covering opportunities for Mercury, outer planets and Mars, contour charts, etc

This paper provides a set of requirements for the technology development of a solar sail propelled Interstellar Heliopause Probe mission. The mission is placed in the context of other outer solar systems missions, ranging from a Kuiper Belt mission through to an Oort cloud mission. Mission requirements are defined and a detailed parametric trajectory analysis and launch date scan performed. Through analysis of the complete mission trade space a set of critical technology development requirements are identified which include an advanced lightweight composite High-Gain Antenna, a high-efficiency Ka-band travelling-wave tube amplifier and a radioisotope thermoelectric generator with power density of approximately 12 W/kg. It is also shown that the Interstellar Heliopause Probe mission necessitates the use of a spinning sail, limiting the direct application of current hardware development activities. A Kuiper Belt mission is then considered as a pre-curser to the Interstellar Heliopause Probe, while it is also shown through study of an Oort cloud mission that the Interstellar Heliopause Probe mission is the likely end-goal of any future solar sail technology development program. As such, the technology requirements identified to enable the Interstellar Heliopause Probe must be enabled through all prior missions, with each mission acting as an enabling facilitator towards the next.

It has been known for some time that Taylor series (TS) integration is among the most efficient and accurate numerical methods in solving differential equations. However, the full benefit of the method has yet to be realized in calculating spacecraft trajectories, for two main reasons. First, most applications of Taylor series to trajectory propagation have focused on relatively simple problems of orbital motion or on specific problems and have not provided general applicability. Second, applications that have been more general have required use of a preprocessor, which inevitably imposes constraints on computational efficiency. The latter approach includes the work of Berryman et al., who solved the planetary n-body problem with relativistic effects. Their work specifically noted the computational inefficiencies arising from use of a preprocessor and pointed out the potential benefit of manually coding derivative routines. In this Engineering Note, we report on a systematic effort to directly implement Taylor series integration in an operational trajectory propagation code: the Spacecraft N-Body Analysis Program (SNAP). The present Taylor series implementation is unique in that it applies to spacecraft virtually anywhere in the solar system and can be used interchangeably with another integration method. SNAP is a high-fidelity trajectory propagator that includes force models for central body gravitation with N X N harmonics, other body gravitation with N X N harmonics, solar radiation pressure, atmospheric drag (for Earth orbits), and spacecraft thrusting (including shadowing). The governing equations are solved using an eighth-order Runge-Kutta Fehlberg (RKF) single-step method with variable step size control. In the present effort, TS is implemented by way of highly integrated subroutines that can be used interchangeably with RKF. This makes it possible to turn TS on or off during various phases of a mission. Current TS force models include central body gravitation with the J2 spherical harmonic, other body gravitation, thrust, constant atmospheric drag from Earth's atmosphere, and solar radiation pressure for a sphere under constant illumination. The purpose of this Engineering Note is to demonstrate the performance of TS integration in an operational trajectory analysis code and to compare it with a standard method, eighth-order RKF. Results show that TS is 16.6 times faster on average and is more accurate in 87.5% of the cases presented.

DOI: 10.2514/1.27610 At the end of its mission to the International Space Station, during its reentry into Earth atmosphere, the automated transfer vehicle is subject to high heat fluxes leading to structural heating and fragmentation of the vehicle. It has been concluded that, depending on the mode of release, onboard residual hypergolic propellants may ignite and explode upon exposure to the hot and reactive flow environment. Because an earlier explosion of the vehicle would change drastically the impact footprint of its fragments onto the Earth surface, this study proposes a reassessment of the explosion potential. From the trajectory analysis, several points of the reentry path have been computed using a Navier–Stokes solver accounting for nonequilibrium effects. Numerical simulations have been performedwithandwithoutperforationofthestructure.Inparallel,acomprehensiveliteraturesurveyonignitionof monomethyl hydrazine and dimethyl hydrazine vapors with pure air or air mixed with nitrogen tetroxide has been performedtoassesstheautoignitionpotentialofthemixture.Finally,theresultsofthecomputational fluiddynamics computationshavebeenusedtoestimatetheexplosionriskinthepresenceofapropellantleakage.Analysisconfirms therisk ofadestruction ofthe automated transfervehicle athigher altitude,which couldinduce adifferent footprint of the fragments on the ground.

The Mars 2001 Odyssey Orbiter successfully completed the aerobraking phase of its mission on 11 January 2002. The support provided by NASA’s Langley Research Center to the navigation team at the Jet Propulsion Laboratory, California Institute of Technology, in the planning and operational support of Mars Odyssey aerobraking is discussed. Specifically, the development of a three-degree-of-freedom aerobraking trajectory simulation and its application to both preflight planning activities and operations is described. The importance of running the simulation in a Monte Carlo fashion to capture the effects of mission and atmospheric uncertainties is demonstrated, and the utility of including predictive logic within the simulation that could mimic operational maneuver decision making is shown. A description is also provided of how the simulation was adapted to support flight operations as both a validation and risk reduction tool and as a means of obtaining a statistical basis for maneuver strategy decisions. This latter application was the first use of Monte Carlo trajectory analysis in an aerobraking mission.

The Mars Pathfinder mission provides the next opportunity for scientific exploration of the surface of Mars following a 7.6 km/s direct entry. In support of this effort, a six-degree-of-freedom trajectory analysis and aerodynamic characteristic assessment are performed to demonstrate vehicle flyability and to quantify the effect that each of numerous uncertainties has upon the nominal mission profile. The entry vehicle is shown to be aerodynamically stable over a large portion of its atmospheric flight. Two low angle-of-attack static instabilities (freestream velocities of about 6.5 and 3.5 km/s) and a low angle-of-attack dynamic instability (supersonic) are identified and shown to cause bounded increases in vehicle attitude. The effects of center-of-gravity placement, entry attitude, vehicle roll rate, aerodynamic misprediction, and atmospheric uncertainty on the vehicle attitude profile and parachute deployment conditions are quantified. A Monte Carlo analysis is performed to statistically assess the combined impact of multiple off-nominal conditions on the nominal flight characteristics. These results suggest that there is a 99.7% probability that the peak attitude throughout the entry will be less than 8.5 deg, the peak heating attitude will be below 6.2 deg, and the attitude at parachute deployment will be less than 3.9 deg.

TRANSYS (TRANsportation SYStem) is a multidisciplinary modularly structured, interactive software system for preliminary design, analysis, and evaluation for a broad range of future space transportation systems. Modules being integrated in the present version, TRANSYS 1.1, are aerodynamics, propulsion, trajectory analysis, geometry, mass estimation, and cost and economics. Each module consists of several application programs and databases. To provide the necessary high degree of flexibility, each module can be applied individually. In the scope of an iterative vehicle design and optimization process, a group of selectable modules has to be executed successively. The data management and the interaction between the individual modules is controlled by an engineering database system. Expendable and reusable launch vehicles with rocket propulsion as well as orbital transfer vehicles can be designed, analyzed, or evaluated with the present version, TRANSYS 1.1.

The conceptual design of a rocket-powered, two-stage, fully reusable launch vehicle has been performed as a part of the NASA's Advanced Manned Launch System study. This paper summarizes the structural design and load analysis of this vehicle. The method used to determine the structural weights consists of generating a finite-element model for each vehicle, selecting a set of critical loading conditions, determining the loads on the model caused by those conditions, determining the model response, and changing the sizes of individual elements to obtain a safe structure. The integrated loads on the vehicle were obtained from a three-degree-of-freedom trajectory analysis.

Electrical System An eight-channel receiver is used to control the functions of the flight vehicle. Of the eight channels, four are assigned to the autopilot and primary flight controls. Servomechanisms are used as the actuating mechanisms. During autopilot operation the three primary flight control channels (elevator, rudder, and ailerons) from the receiver are disabled, allowing signals from only the autopilot to activate the servomechanisms. Conversely, when the autopilot is disengaged, the three flight control channels are active. This approach allows only one set of signals to reach the flight controls, either from the autopilot or from the ground controller. The remaining four channels are assigned to engine start, wing lock release, landing-gear release, and flaps. At the instant of engine start, a timer is activated, and the wing lock is released at a predetermined elapsed time. This serves as a backup feature, allowing wing deployment to occur in the event that no signals are received from the ground controller. Trajectory Analysis An analysis of the trajectory was conducted with SORT.2 This program allows three-dimensional trajectory optimization. Only pointmass trajectories were analyzed. Engine burnout occurs 36 s after liftoff, at an altitude of 12,965 ft. The maximum altitude achieved by the vehicle is 18,870 ft. Concluding Remarks This paper indicates that it may be possible to design a simple reusable rocket. The next step will be to build and test such a vehicle. This design could be extended and adapted to larger vehicles in order to attain the higher altitudes that are required in some of the applications of sounding rockets, such as upper-atmosph eric experiments.

The aerodynamic database upon which the missions and the proposed designs of an atmospheric entry research vehicle (ERV) would be based are discussed, along with a trajectory analysis of two missions proposed for the ERV. The ERV is intended as a means to explore reentry techniques, thermal protection systems and cross-range and orbital change maneuvers beyond the operational envelope of the Orbiter. The 25 ft long ERV would have a 13 ft wingspan, allowing it to be launched from the Orbiter bay. The ERV would have atmospheric maneuvering control surfaces and reaction control systems for orbital maneuvers. The total weight would be 12,000 lb. The aerodynamic design for the ERV was developed from wind tunnel, theoretical and Shuttle data for continuum, transition and free-molecule flight regimes. Typical flight profiles for intended missions are outlined.

The potential performance advantages of solar electric propulsion have led to a study of the over-all spacecraft systems implications of such propulsion devices. The results of this study, including spacecraft design and trajectory analysis (using the Atlas/Centaur as the launch vehicle) for asteroid belt, out-of-ecliptic, and solar probes are presented. These design studies have investigated the feasibility of a baseline spacecraft that could accommodate the requirements and constraints of each mission considered and indicate the degree of subsystem flexibility required of such a multipurpose spacecraft. Particular attention is devoted to a modular-ion engine system design that can perform a variety of missions with no basic modifications to major elements. A fixed attitude spacecraft design able to perform efficiently the three selected missions is described.

2 Hitt, E. F. and Rea, F. G., "Development of an Evaluation Technique for Strapdown Guidance Systems/' Second Interim Scientific Report, Contract NAS 12-550, Feb. 28, 1969} BatteUe Memorial Institute, Columbus Labs., Columbus, Ohio. s Rea, F. G. and Fischer, N. H., "An Improved Method of Estimating Midcourse Fuel Requirements (Approximating the Probability Distribution of the Magnitude of a Vector with Normal, Zero Mean, Components)," Paper presented to NASA/ ERG Fourth Guidance Theory and Trajectory Analysis Seminar, Cambridge, Mass., May 16-17, 1968, *Strack, W. C. and Huff, V. N., "The N-Body Code, A General FORTRAN Code for the Numerical Solution of Space Mechanics Problems on an IBM 7090 Computer," TN D-1730, Nov. 1963, NASA. 5 "Saturn IB Improvement Study (Solid First Stage) Phase II, Final Detailed Report,'' Kept. SM-51896, Vol. II, March 30, 1966, Douglas Missile and Space Systems Div., Huntington Beach, Calif. 6 "Launch Vehicle Estimating Factors/' Jan. 1969, NASA. 7 "A Study of Jupiter Flyby Missions, Final Technical Report," FZM-4625, May 17, i966, General Dynamics, Fort Worth Div., Fort Worth, Texas.

A trajectory analysis is presented which shows the advantage of lifting ascent and the penalties associated with increased wing mass due to increases in aerodynamic loading. The vehicle analyzed is made up of an orbiter, an expendable liquid fuel tank, and two solid rocket motors. Results of the tradeoff are illustrated in a graph of the total vehicle mass plotted against the trajectory constraint. Results indicate that even when structural mass penalties are considered lifting ascent provides a payload delivery.