Engineering topics
Publications and source records attributed to Kwok, J. H..
This paper describes a nominal mission plan that progressively establishes SIRTF capabilities during the IOC/SV phases, taking into consideration thermal, cryogrenic, optical, communications, celestial mechanics, and operational designs and constraints.
Designed for use on IBM PC, program suited for lifetime analysis of orbiting spacecraft.
Drifts and lifetimes predicted. Long-term Orbit Predictor (LOP) is trajectory-propagation computer program used as analysis tool for studies of lifetimes of orbiting spacecraft. Used for any planetary orbiting missions. LOP written in FORTRAN 77.
This paper describes the implementation of two computer programs that predict the motion of artificial satellites. The two programs are developed in a microcomputer environment. The particular microcomputer used is an IBM AT system. The programs can predict the orbital behavior of artificial satellites under the influence of gravity harmonics up to a 21 by 21 field, luni-solar gravity, drag, and solar radiation pressure. One program uses Cowell's method and the other uses variation of parameters with a singly averaging technique. Both programs use commercially available software to post-process trajectory information and to generate graphics display. The programs reside on floppy disks and are available to the general public through NASA's Computer Software Management and Information Center (COSMIC).
Two families of symmetric periodic orbits of the planar, general, three-body problem are presented. The masses of the three bodies include ratios equal to the sun-Jupiter-Saturn system and the periods of the orbits of Jupiter and Saturn are in a 2:5 resonance. The (linear) stability of the orbits are studied in relation to eccentricity and mass variations. The generation of the two families of periodic orbits follows a systematic approach and employs (numerical) continuation from periodic orbits of the first and second kind in the circular restricted problem to the elliptic restricted problem and from the circular and elliptic problems to the general problem through bifurcation phenomena relating the three dynamical systems. The approach also provides insight into the evolutionary process of periodic orbits continued from the restricted problems to the general problem.
A set of singly averaged equations of motion are presented and applied to long-term orbit prediction of an orbiting spacecraft around a slowly rotating planet, using the Venus Radar Mapper Mission as an example. The equations of motion used are valid for all eccentricities less than one. The disturbing potentials used include nonsphericity of the Venus gravity field and third-body effects due to the sun. Recursive relationships are used in the expansion and evaluation of these potentials and their respective partial derivatives. Special care is taken to optimize computational efficiency. The averaging method is compared with high precision Cowell's method using a desktop microcomputer and shows computational saving of about two orders of magnitude.
The Venus Radar Mapper (VRM) mission is sponsored by NASA to put a single spacecraft in orbit around Venus to map the surface of Venus using a synthetic aperture mapping radar. The spacecraft is scheduled to be launched in April 1988 using a Shuttle-Centaur G combination. The spacecraft arrives at Venus in late July 1988 and begins its mapping mission which lasts for one Venus rotation or 243 days. This paper describes the VRM mission at its present state of design. The science objectives and project constraints are described. Key features of the spacecraft system and radar system are discussed. The interplanetary and mapping orbit design are covered. Navigation strategy is explained, including trajectory maneuvers and mapping phase orbit determination. Finally, the mapping sequences to optimize planet coverage are described.
A conceptual design is outlined for the navigation subsystem of the Autonomous Redundancy and Maintenance Management Subsystem (ARMMS). The principal function of this navigation subsystem is to maintain the spacecraft over a specified equatorial longitude to within + or - 3 deg. In addition, the navigation subsystem must detect and correct internal faults. It comprises elements for a navigation executive and for orbit determination, trajectory, maneuver planning, and maneuver command. Each of these elements is described. The navigation subsystem is to be used in the DSCS III spacecraft.
Results are presented for a mission and trajectory design for a Venus Radar Mapper mission that will place a SAR in a nearly polar orbit around Venus. The mission is intended to obtain images of at least 70% of the planet's surface at a line-pair resolution of 1 km and to produce global maps of the planet's topography and gravity field. The mission design criteria are discussed, along with the spacecraft flight system, the mission performance and domains, a mission profile, the mapping orbit and SAR coverage, and the data acquisition strategy.
Results of mission and trajectory design for the Venus Orbiting Imaging Radar are presented. Because of the application of the aerobraking technique, for the first time to a planetary mission, the conventional mission design criteria in the selection of the launch/arrival dates must be modified and new mission constraints must be imposed. Venus itself poses an interesting aspect in trajectory design for an orbiting spacecraft due to its dense lower atmosphere and its slow retrograde rotation. Some recent results in the method of averaging applied to an orbiting spacecraft about this unique planet are given.
A systematic approach to generate periodic orbits in the elliptic restricted problem of three bodies is introduced. The approach is based on (numerical) continuation from periodic orbits of the first and second kind in the circular restricted problem to periodic orbits in the elliptic restricted problem. Two families of periodic orbits of the elliptic restricted problem are found by this approach. The mass ratio of the primaries of these orbits is equal to that of the Sun-Jupiter system. The sidereal mean motions between the infinitesimal body and the smaller primary are in a 2:5 resonance, so as to approximate the Sun-Jupiter-Saturn system. The lineaar stabilities of these periodic orbits are studied as functions of the eccentricities of the primaries and of the infinitesimal body. The results show that both stable and unstable periodic orbits exist in the elliptic restricted problem that are close to the actual Sun-Jupiter-Saturn system. However, the periodic orbit closest to the actual Sun-Jupiter-Saturn system is (linearly) stable.
Aerobraking techniques are applied to trajectory design for the VOIR (Venus Orbiting Imaging Radar) mission to enhance overall mission performance. The approach can be used for other planetary missions such as a Titan orbiter or a Mars sample return to earth. The orientation of the elliptical insertion orbit must be chosen in such a way that the combined effects of solar gravity, central-body harmonics, and drag on periapsis altitude are minimized to avoid frequent maneuvers. The selection of the base and width of a corridor for periapsis altitude depends upon the temperature and integrated heat of the aerobrake shield, the atmospheric density profile, the duration of aerobraking, the number of maneuvers, the time available between maneuvers, and uncertainties in the central-body gravity field and atmospheric density. Flight path profiles for an aerodynamically stable spacecraft and an inertially fixed spacecraft passing through the free molecular flow regime, as well as the transitional flow regime, are illustrated
The scientific objectives and rationale for a Venus Orbiting Imaging Radar Mission are presented. A provisional science payload responsive to these objectives is described and a reference set of measurement requirements and their priorities is established. Those high priority measurements that are most demanding on spacecraft and mission design are used to develop a reference spacecraft design and a reference mission design. A discussion of mission performance issues is also included.
In a paper by the second author (Nacozy, 1981), various time elements are presented for use with the Sundman time transformation. In that paper, the time elements are given in terms of Keplerian orbital elements. We give here the corresponding time elements in terms of rectangular coordinates. Extensive references are given in the previous paper and will be omitted here. We present additional numerical experiments comparing the use of time elements and time transformations together with the use of time transformations alone. The results indicate a reduction in computational error when time elements are used.
Periodic orbits of the asteroidal type in the circular restricted problem are studied by varying the period of the infinitesimal body (asteroid) and the mass ratio of the primaries (Sun-Jupiter mass ratio). The results indicate that asteroidal periodic orbits can exist for the actual Sun-Jupiter mass ratio for resonances of the Hecuba (2:1), Hilda (3:2) and Thule (4:3) groups, but not for resonances of higher consecutive integer ratios. It is also found that an asteroid can be placed in a periodic orbit at a position with mean motion between 2:1 and 3:2 even if Jupiter is about 9 times more massive than its actual value. However, an asteroid cannot be placed in a periodic orbit beyond the 4:3 resonance for the actual Sun-Jupiter mass ratio.