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This paper describes an operational interactive computer graphics system, called Graphic ROCKET, currently being used as a performance analysis aid for preliminary design and evaluation of space vehicle systems. It provides means for rapidly specifying and evaluating the performance of a wide range aerospace vehicle designs and flight plans.

The report was written to provide a sample problem to be used to guide the user of the water droplet impingement prediction computer program. The report is intended to illustrate the preparation of input data and the interpretation of results.

Trajectory optimization, performance and other factors in analysis of low thrust solar electric propulsion Jupiter flyby mission

Mars lander deorbit trajectory sensitivity analysis for fixed flight path and communications angles at atmospheric entry, comparing with Monte Carlo simulation

Analysis of the boost-phase performance of a rocketboosted missile generally requires the solution of a set of differential equations with variable coefficients. Analytical methods are difficult and time consuming if not impossi­ ble. This paper discusses four numerical approaches to the solution of the equations for a cruciform, uncontrolled missile booster. The first of these, called the "exact" method, will yield results which are as accurate as one cares to make them. The second, referred to as the ""semiexact," assumes that deviations from the flight path are small and oscillatory. With this assumption the solution may be completed in two parts: (a) Computation of the speed-time history and (b) substitution of the speed into the remaining equations. The accuracy for short time oblique launch or for long time vertical launch is as good as the "exact." The third, applicable under the same conditions as the "semi-exact" and referred to as "linear speed-time history," assumes a constant acceleration from launch to the end of boost. The accuracy of this method is very good (on the order of 2 per cent for the terminal speed) provided the thrust-time diagram is nearly trape­ zoidal. The fourth, called "iterative solution for separa­ tion speed," yields an approximate value for the terminal speed with accuracy on the order of 2-4 per cent. The various uses of the different methods are also discussed as well as techniques for including cross winds, estimating separation speeds and tolerances allowed in estimation of step-wise acceleration. Sample calculations for the three approximate methods are given.

The TRACE Trajectory Analysis and Orbit Determination Program is a general-purpose orbital analysis program. It was written and continues to be developed specifically to assist technical personnel in the analysis and design of satellite orbits and tracking systems. Volume VII is intended to serve as in input usage guide that defines all input required to perform TRACE functions such as trajectory generation, data/observation generation, orbit determination, and statistical analysis. The report contains the appendices which includes: Description of the general-purpose input routine GAIL 1; Trace data deck structure examples; Sample output descriptions; and Trace input forms.

The TRACE Trajectory Analysis and Orbit Determination Program is a general-purpose orbital analysis program. It was written and continues to be developed specifically to assist technical personnel in the analysis and design of satellite orbits and tracking systems. Volume 7 is intended to serve as in input usage guide that defines all input required to perform TRACE functions such as trajectory generation, data/observation generation, orbit determination, and statistical analysis.

The Trajectory Analysis Program (TAP) is designed specifically for use on the CDC 6000 series machine as a computational tool to generate trajectories by different mathematical formulations and to compare the results. In particular, the reference trajectory is the result of numerical integration of the equations of motion in a Cowell formulation with time as the independent variable; whereas, the comparison trajectory selected is achieved by an analytic (Pines, Kyner, or Escobal) or a numerically integrated Encke (classical or modified) formulation. A comprehensive description is presented of the algorithms implemented in TAP with emphasis on characteristics such as logic flow, equations, numerical techniques, comparison differencing, and plotting. A basic usage guide is included with complete instructions for input data preparation and explanations of output tables.

This paper summarizes a portion of the planetary mission analysis results of past and present studies conducted by Rockwell International for NASA-MSFC (Contract NAS8-27360) dealing with the feasibility of a Solar Electric Propulsion Stage (SEPS). The SEPS is envisioned as an upper stage of a transportation system capable of delivering either separable payload spacecraft or attached science packages to various planetary targets. The purpose of the paper is to demonstrate that, from a payload performance capability standpoint, a common SEP Stage can deliver various payloads to a host of planetary targets including inner and outer planets, asteroids, and comets.

Methodology employed in development of a computer program designed to analyze optimal low-thrust trajectories is described, and application of the program to a Solar Electric Propulsion Stage (SEPS) geosynchronous mission is discussed. To avoid the zero inclination and eccentricity singularities which plague many small-force perturbation techniques, a special set of state variables (equinoctial) is used. Adjoint equations are derived for the minimum time problem and are also free from the singularities. Solutions to the state and adjoint equations are obtained by both orbit averaging and precision numerical integration; an evaluation of these approaches is made.

A summary of the performance of Solar Electric Propulsion spacecraft for Comet Encke missions for the 1980, 1984 and 1987 mission opportunities is presented together with a description of the spacecraft trajectory for each opportunity. Included is data for rendezvous trajectories for all three opportunities and data for a slow flyby mission during the 1980 opportunity. A range of propulsion system input powers of 10 to 20 kW are considered together with a constant spacecraft power requirement of 400 watts. The performance presented in this paper is indicative of that using 30 cm Mercury electron bombardment thrusters that are currently being developed. Performance is given in terms of final spacecraft mass and is thus independent of any particular spacecraft design concept.

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

A spark chamber track finding program has been written. The accuracy of the location of the track is good, and if the chamber contains many spurious sparks the computation time is much less than the time used by a least squares fit routine. Multiple tracks can be found and interaction vertices can thereby be determined.

The magnetic coordinates of the Pioneer 10 spacecraft and the five innermost satellites are reported for the Jupiter encounter. The D sub 2 offset is used to make the calculations. Magnetic coordinates are needed for the interpretation of the trapped particle measurements, including the absorption effects of the satellites. Contours of constant field magnitude and magnetic latitude are given at the surface of Jupiter for the D sub 2 model. The system 3 longitude of a spacecraft at Jupiter is derived, and formulas given for the relationships between system 1, 2, and 3 longitudes. The longitude of the magnetic dipole increases by about 3 deg per year, due to the inaccurate rotation rate used to define system 3 longitude.

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

Navigation data from Mariner Mars 1969 TV pictures, estimating spacecraft trajectory to planetary center

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

Launch vehicle injection error sensitivity minimization by trajectory shaping using excess booster capability

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

A review of engineering and research data requirements for altitudes 30-100 km is made, indicating a variety of concrete data applications above 30 km (10 mb). The required data have in large measure been provided from meteorological rocket soundings, of which 18,000 have been taken, at considerable cost, since 1959. Remote vertical soundings based on satellite infrared radiation measurements have been obtained since 1969 with considerable success to an altitude of 25 km (30 mb). From developmental work in progress, it is expected that reliable temperature soundings may be obtained to 40-45 km. A discussion of the overall reliability and utility of the satellite data leads to several conclusions regarding the continuing need for rocket soundings.

Satellite-star optical navigation, a new spacecraft navigation technique has been successfully demonstrated during Mariner 9's approach to Mars in mid-November. As Mariner 9 approached Mars, 21 TV pictures of its satellites, Phobos and Deimos, were taken and transmitted to earth. On earth, these pictures were carefully processed in order to locate the positions of the bright satellites against the relatively faint star background. The objectives and performance of the preflight and real time activities leading to the successful demonstration of satellite-star navigation data are described.

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.

Linearized manual space navigation incorporating redundant measurements to compensate for instrumentation inaccuracies

Computer program for current profile and trajectory analysis for cesium contact thrustor leading to integral focus configuration of multistrip contact engines for long space missions

Heliocentric low thrust spacecraft trajectory analysis using two-variable asymptotic expansion method

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

Optimum launch trajectories for ATS-E mission with noncircular parking orbits, considering apogee motor size, perigee radius and duration constraints

Wind tunnel free-flight testing of reentry vehicle dynamics, noting capability for simultaneous two- dimensional viewing of model trajectories

Manned space flight safety through mission trajectory design, considering hardware, software and operational constraints

Matched asymptotic expansion method applied to hypervelocity flight trajectories, obtaining solutions to flight dynamic equations

N body numerical integrator using virtual mass technique for computing trajectories with speed and accuracy suitable for spacecraft onboard computer applications

Translational forces on Mariner 5 stemming from attitude control system studied to determine trajectory for scientific and operational purposes

Postflight analysis including six degree of freedom trajectory digital simulation of Aerobee 350 sounding rocket behavior under large thrust misalignment

A rapid and accurate method of calculating optimal impulsive transfers in the restricted problem of three bodies has been developed. The technique combines a multi-conic method of trajectory integration with primer vector theory and an accelerated gradient method of trajectory optimization. A unique feature is that the state transition matrix and the primer vector are found analytical without additional integrations or differentiations. The method has been applied to the determination of optimal two and three impulse transfers between the L2 libration point and circular orbits about both the earth and the moon.

Pioneer-F Jupiter flyby mission planning, considering use of optimum broken plane trajectories to increase launch opportunity and/or payload capability

Trajectory design and mission planning for Mariner Venus 67 /Mariner 5/ space project, discussing perturbative forces, orbit determination and midcourse guidance

Comparative sting supported and free flight tests in hypersonic wind tunnel on modified Apollo launch configuration

The application of digital computers to reduction of telemetry data assumes an increasingly important role in the analysis of physical problems. Even after reduction of the raw sensor data, an immense volume of resultant data remains. Thus there is a constant need of new tools for computer-aided analysis, synthesis and display of scientific results.

Trajectory characteristics of high-thrust Mars missions and effect on other aspects of mission planning

Parachute trajectory and opening load prediction based on inflation process and added mass, determining drag area as function of distance