Optimal aeroassisted return from high earth orbit with plane change

This paper gives a complete analysis of the problem of aeroassisted return from a high Earth orbit to a low Earth orbit with plane change. A discussion of pure propulsive maneuver leads to the necessary change for improvement of the fuel consumption by inserting in the middle of the trajectory an atmospheric phase to obtain all or part of the required plane change. The variational problem is reduced to a parametric optimization problem by using the known results in optimal impulsive transfer and solving the atmospheric turning problem for storage and use in the optimization process. The coupling effect between space maneuver and atmospheric maneuver is discussed. Depending on the values of the plane change i, the ratios of the radii, $\text{n =}\text{r}{}_1\text{r}{}_2$ between the orbits and $\text{a =}\text{r}{}_2\text{R}$ between the low orbit and the atmosphere, and the maximum lift-to-drag ratio E1 of the vehicle, the optimal maneuver can be pure propulsive or aeroassisted. For aeroassisted maneuver, the optimal mode can be parabolic, which requires only drag capability of the vehicle, or elliptic. In the elliptic mode, it can be by one-impulse for deorbit and one or two-impulse in postatmospheric flight, or by two-impulse for deorbit with only one impulse for final circularization. It is shown that whenever an impulse is applied, a plane change is made. The necessary conditions for the optimal split of the plane changes are derived and mechanized in a program routine for obtaining the solution.     

Communications transponder performance for Japanese Communications Satellite-2 and prospects for future systems

Communications transponder for the Japanese Communications Satellite-2 (CS-2a and 2b) to be launched into a geostationary orbit by N-II launch vehicle in February and August, 1983, has been developed. The transponder is provided with six-channel K-band ($\text{30}\text{20}\text{GHz}$) transponder including beacon transmitter, which operates in the highest frequency ranges ever utilized on an operational communications satellite, and two-channel C-band ($\text{6}\text{4}\text{GHz}$) transponder. Receiver front end of the K-band transponder consists of a direct mixer followed by a 1.8 GHz IF amplifier and provides 8 dB noise figure. 20 GHz output power is 4 W by final amplification at 5-W TWTA. C-band transponder provides 4 dB noise figure and 4.3-W output power. Key factors for future high capacity transponder are also presented.     

A small parameter method in problems of maneuvering space vehicles with aerodynamic efficiency

The motion of a satellite with aerodynamic efficiency along a low near-circular orbit is considered in the paper. The controls of bank angle γ and lift coefficient C_y are used as control functions. The introduction of a small parameter ($\text{? = (}\text{?}{}_0\text{· S · g}{}_0\text{2G}\text{)}$) makes it possible to integrate an adjoint system of equations and to obtain an approximate solution to the complete problem in the class of piecewise-constant control functions. Maximum values for the coordinates of heading angle η and lateral derivation from the plane of a reference orbit ?, which are connected with orbit plane angle by the relation cos i = cos ? · cos h, are used as criteria of maneuvering capability for a satellite with aerodynamic efficiency. Optimal programs for bank angle and incidence variation are derived and the influence of lift-to-drag ratio on the vehicle maneuvering capabilities has been estimated.It is shown that the process of the optimal motion is a special kind of gravitational skipping similar to the Keplerian motion but with continuous descent.     

Results of the early orbit manoeuvres of the ampte-uks

The Active Magnetospheric Particle Tracer Explorers (AMPTE) program consists of three satellites which were launched on 16th August 1984. The scientific aim of the mission is to inject lithium and barium tracer ions inside and outside the Earth's magnetosphere and to detect and monitor these ions as they diffuse through the inner magnetosphere. The first of these three satellites, the U.S. Charge Composition Explorer (CCE) was launched into an elliptical orbit of apogee 8 R_e. The other two satellites are the West German Ion Release Module (IRM) and the U.K. Subsatellite (UKS), both of which were launched on the same vehicle into a highly elliptical orbit of apogee 18 R_e. At discreet intervals during the mission the IRM will release ions into the solar wind, and the movement of these ions will be monitored by the UKS. Depending on the particular scientific requirement, the UKS has to be positioned accurately at a given distance behind the IRM. Initially the UKS has to be located 100 km behind the IRM, and held there for ～9 months. It will then be moved a distance of ～1 R_e behind the IRM. In order to manoeuvre the UKS around its orbit, a cold gas jet system is incorporated on the satellite, allowing impulses to be applied both along and perpendicular to the orbit velocity vector. The orbit control system also has to cater for relative orbit changes due to air drag at perigee, as the IRM and the UKS have different $\text{area}\text{mass}$ ratios. This paper presents an account of the orbit control system implemented on the UKS, together with the mathematical approach adopted, and results from manoeuvres made in the first weeks of the mission.     

On orbit control using gravity gradient effects

Possibility of orbit control using gravity gradient (GG) effects without any mass expulsion is discussed. For simplicity, a dumb-bell type satellite and circular orbits are mainly considered. It is shown that the GG effects can be applied to convert attitude torques into orbital torques and vice versa. In central gravitational force fields, maximum orbital torques or thrusts are available from the GG force when roll or pitch angle is $\text{±}\text{π}\text{4}$ provided that the attitude angle is null when the dumb-bell axis coincides with the local vertical. Such external torques as geomagnetic or solar wind pressure can be utilized to maintain the $\text{±}\text{π}\text{4}$ attitude, then the orbital torques are available forever. In non-central gravitational fields, without any external torque, the orbital radii of circular orbits can be increased by controlling the satellite attitude using electric energy. The use of the Earth's oblateness effects and the exterior Lunar potential is discussed.     

A practical approach to the disposal of highly toxic and long-lived spent nuclear fuel waste between Venus and Earth

Extraterrestrial disposal, while not the only alternative (there is at least one very safe terrestrial method), nevertheless assures definite and irreversible removal of the most toxic and long-lived waste from the biosphere. In the foreseeable future, there is far less need to retrieve nuclear waste for later use then to dispose of it in a publicly acceptable manner, beginning in the near future (middle 1980s). It is, therefore, more important to assure safety in the weakest link of this disposal process—transportation into near-Earth orbit—than to engage in a retrievable disposal beyond Earth. The disposal “site” should lie at minimum safe transfer energy level. Primary candidate is the space between Venus and Earth. The number of propulsion phases should be a minimum, preferably only one (out of near-Earth orbit). Lunar gravity assist can be helpful to achieve higher inclination of the heliocentric orbit relative to the ecliptic.Solidified spent fuel isotopes and actinides, sufficient to reduce the residual terrestrial waste to the radiation level of natural uranium deposits after 30–40 yr instead of 1000–1500 yr, is deposited into heliocentric orbits. Transportation systems, requirements, costs and the associated socio-economic benefit potentials of an environmentally more benign and a more vigorous nuclear power generation program are presented.Prior to solidification, an interim storage of 10 yr, following removal from the reactor, may be required. The Shuttle, with one Orbiter modified as Nuclear Waste Carrying Orbiter (NWCO) and an out of near-Earth orbit booster, provides a safe and economic transportation system at (1979$) disposal mission costs from surface to disposal orbit of less than $\text{0.5¢/}\text{kWhe or}\text{? 0.1¢/}\text{kWhe}$ (some 70% of overall exo-disposal cost), depending on level of orbital operations (this at basic Shuttle flight cost of $30M). The orbital operations mode can be selected according to lead time and costs, and can be advanced sequentially, lowering disposal costs while at the same time financing the development of orbital operations techniques needed also for other and larger-scale exo-industrial activities. An average of 10–15 disposal missions of the NWCO is required annually, at the given conditions, to service the spent fuel of 173 reactors 1978 in operation in W. Europe, Japan and North America.     

Recovery of long-wavelength mean gravity anomalies for the dedicated gravitational satellite (gravsat) mission

This paper describes a computer simulation study that was undertaken to determine how well long-wavelength variations in the Earth's gravitational field can be recovered using data from the Dedicated$\text{Grav}$itational $\text{Sat}$ellite (GRAVSAT) mission. This mission is to consist of two low altitude (160 km) spacecraft in essentially the same orbit but separated in phase by 100–300 km. Geodetic data are measurements of the relative range rate to an accuracy of about 1 μm/s at 4 sec intervals. Specifically, a Bayesian covariance simulation was used to investigate simultaneous recovery of the spacecraft ephemerides and a global distribution of 20° × 20° mean gravity anomaly blocks. Sources of errors considered were tracking station positions, gravitational constant, Earth body tides, tropospheric modeling and measurement noise. It should be noted that this simulation does not include as an error source variations in the gravity field that have a character different from what was modeled. Consequently, this study demonstrates the potential of the low-low system as configured to recover the long-wavelength variations in the gravity field.Using only one days worth of data, the mean of the standard deviations of the 162 20° × 20° gravity anomaly blocks is about 1 μgal. For a 6 month mission (assuming a reduction proportional to the square root of the data intervals) this projects to < 0.1 μgal. Because of the potential of increased measurement precision at shorter separation distances, and the relative insensitivity of the recovery process to separation distance, it should be possible to recover both long and short wavelength variations with a modest distribution of separation distances tailored primarily to the short wavelength recovery. Effects of the uncertainty in the gravitational constant and Love's numbers are negligible. In a simulation not reported on, increasing the altitude of the orbit to 200 km from 150 km, degraded, as expected, the accuracy of the recovered parameters by only 7%.     

Harenodynamic cooling: The use of lunar sand as cooling medium

In the context of investigations to maximize the operational utilization of local resources on other worlds, a concept for using fine lunar sand as cooling medium for the rejection of waste heat is developed. Sand at cryogenic temperature is admixed to the working fluid containing the waste heat. Gas and sand form a mix of the desired and temperature of the working fluid. Subsequently, the sand is removed and intensely radiation-cooled, utilizing its very large $\text{surface}\text{volume}$ ratio. Thereafter it is recycled. It is shown that, compared to heat rejection by radiator, harenodynamic cooling offers superior efficiency, greater compactness and lower cost, improving the economics of nuclear power plants and of industrial expansion on the Moon in general.     

火星探测器使命轨道捕获策略研究

分析了火星探测器使命轨道捕获策略研究需求,给出了三种可用的典型捕获策略,建立了速度纵向平面内的飞行动力学方程,对三种典型策略进行了飞行动力学仿真,给出了各策略完成捕获所需速度增量、任务耗时及过。通过对仿真结果的分析比较,得出研究结论为：直接制动捕获任务耗时及过载较小,但速度增量需求较大;多次穿越后捕获任务耗时较长,但速度增量需求和过载较小;一次穿越后捕获速度增量需求较小且任务耗时较短,但过载较大。     

Equilibrium positions of a weight on a cable fixed to a dumbbell-shaped space station moving along a circular geocentric orbit

The motion of a space object in the gravitational field of the Earth is considered. The object consists of an extended space station and a weight, which is free to move along the cable fixed to the ends of the station. It is assumed that the station is composed of two masses coupled by a weightless rod, while the cable is weightless and non-stretched. The equations of motion of such a system are derived for the case when the motion proceeds in a single plane, while the center of mass of the system moves along a circular geocentric orbit. The conditions of the cable tension (conditions of being on tie) are derived. The phase portrait of the weight motion along the cable is constructed when the station is oriented to the attracting center or is perpendicular to this position. The possibility to leave the tie in this case is analyzed. Equilibrium configurations of the system are found, i.e., such motions of the object under consideration at which the weight does not change its position relative to the station. Lyapunov stability of such configurations is analyzed for two situations: when the station is composed of equal masses and when masses at the ends of the station are different. In particular, for the case of different masses it is established that there exist such positions of equilibrium at which the dumbbell is located at an angle to the direction to the attracting center. In some cases these positions can be stabilized (if the weight is fixed on the cable).     

Collisions of stars by oscillating orbits of the second kind

Ordinary estimations of the number of star collisions in our galaxy—by simple kinematic considerations—lead to a very small number of such collisions: about one or even less every millions of years. However star collisions can occur through the following indirect way which has a much higher probability. (a) Binary stars are very common in our galaxy, about 30–50% of the stars. (b) If two binary stars meet a triple system can be formed by an ordinary exchange type motion. (c) A triple system is generally decomposed into the “inner orbit” (i.e. the relative orbit of the two nearest stars) and the “outer orbit” (i.e. the relative orbit of the third star with respect to the center of mass of the two nearest stars). The major axes of these two orbits have generally small perturbations and it is the same for the eccentricity of the outer orbit. On the contrary, if the relative inclination of the two orbits is large, the perturbations of the eccentricity of the inner orbit are important and can even in some cases lead to an eccentricity equal to one, that is to a collision of the two stars of the inner orbit.Such orbits can be called “oscillating orbits of the second kind”, indeed the first oscillating orbits—conceived by Khilmi and described for the first time in an example by Sitnikov—have unbounded mutual distances r_ij, but the system always come back to small sizes, it has an infinite number of very large expansions followed by strong contractions and, in the three-body case, an upper bound of lim inf (r_1.2 + r_1.3 + r_2.3) can be given in terms of the three masses and the integrals of motion. For the oscillating orbits of the second kind the mutual distances r_ij are bounded, but the velocities are unbounded (i.e. lim inf r_ij = 0 for at least one r_ij) and the system goes to a collision if the bodies have non-zero radius even small. The analytical study of the oscillating orbits of the second kind is a part of the general analytical study of the three-body problem, a part which must be valid for large eccentricities and large inclinations. The use of Delaunay's variables and of a Von Zeipel transformation lead to a first order integrable approximation, valid for any eccentricities and any inclinations, and giving the following results: (a) The oscillating orbits of the second kind occur when the angular momentum of the outer orbit has a modulus sufficiently close to the modulus of the total angular momentum of the three-body system. Hence these orbits occur for inclinations in the vicinity of 90°. (b) The oscillating orbits represent a set of positive measure of phase space and the first order study allows to give a rough estimation of the probability of collisions—even for stars of infinitely small radius. This probability, for given initial major axes and eccentricities and for isotropic arbitrary initial orientations, is generally of the order of m₃$\text{R}\text{M}$ (m₃ being the mass of the outer star, M the total mass and R the ratio of the period of the inner orbit to that of outer orbit).One question remains to be solved: how many collisions of stars are due to that phenomenon? That question is difficult because the probability of formation of a triple system by a random meeting of two binaries is very uneasy to estimate. However it seems that, compared to the usual evaluations based on pure kinematic considerations without gravitational effects, the number of collisions must be multiplied by a factor between one thousand and one million.     

Challenges and perspectives of transport cargo vehicles utilization for performing research in free flight

Russian Progress transport cargo vehicles have successfully been used in different space station programs since 1978. At present time, they play an important role in the International Space Station (ISS) project. Main tasks performed by the transport cargo vehicle (TCV) in the station program are the following: refueling of the station, delivery of consumables and equipment, waste removal, station attitude control and orbit correction maneuver execution.     

Geostationary station keeping control of a space elevator during initial cable deployment

This paper presents a basic concept to derive an orbital control strategy to achieve the full deployment and the geostationary station keeping of a space elevator during its initial cable deployment. The space elevator model is composed of a main spacecraft, a sinker mass and a massive cable connecting them. The cable elasticity, flexibility and taper of the cross-sectional area are omitted for simplification. A reference trajectory is designed so that the space elevator and its center of mass ascend vertically along the geostationary position with keeping the geostationary orbital rate. From the reference trajectory analyses, an orbital control that leads the space elevator orbit to the reference one is derived. However it is found that the reference trajectory is unstable throughout the deployment and a linear feedback control is introduced for stabilization. It is also clarified that the libration destabilizes the orbital control because the orbital acceleration caused by the libration always acts in the opposite direction to the orbital control. Therefore, a libration control is also introduced to stabilize the coupled orbital and librational motions. Numerical simulation result clearly shows that these controls facilitate the full deployment and the geostationary station keeping of the space elevator within the feasible thrust force and amount of propellant.     

Analytical solutions for the relative motion of spacecraft subject to Lorentz-force perturbations

A spacecraft capable of producing higher-than-natural electrostatic charges may achieve propellantless orbital maneuvering via the Lorentz-force interaction with a planetary magnetic field. Development of maneuver strategies for these propellantless vehicles is complicated by the fact that the perturbative Lorentz force acts along only a single line of action at any instant. Relative-motion dynamical models are developed that lead to approximate analytical solutions for the motion of charged spacecraft subject to the Lorentz force. These solutions indicate that the principal effects of the Lorentz force on a spacecraft in a circular orbit are to change the intrack position and to change the orbit plane. A rendezvous example is presented in which a spacecraft with a specific charge of ?3.81 $\times$ 10^?4 C/kg reaches a target vehicle initially 10 km away (on the same equatorial low-Earth orbit) in 1 day. Fly-around maneuvers may be achieved in low-Earth orbit with specific charges on the order of 0.001 C/kg.     

Implementation of a semianalytic satellite theory with recovery of short period terms

First order averaging is applied to the artificial satellite problem to obtain the averaged orbit which includes the secular, long and medium period effects of the oblateness of the Earth and the third body perturbations of the moon and sun. Perturbation theory is then used to recover the short period effects due to J₂, the moon, and sun. The perturbation analysis is carried out by means of Lie series and is developed through the first order. Optimization of the resulting short period series was then accomplished in several steps: first all separate algebraic coefficients were precalculated and stored; then all redundant SIN/COS calls were eliminated; next all repetition of numeric and algebraic coefficients were precalculated in pairs; application of the distributive principle allowed a significant reduction in additions and multiplications; finally trigonometric identities were used to further reduce the SIN/COS computations. The result of this optimization along with an interpolator for the averaged equations of motion results in a computer program which requires only $\text{1}\text{6}$ the CPU time (with no loss in accuracy) of the original non-optimized test program.     

Orbit prediction for IRAS using vector and analytic techniques

This paper describes the techniques of a vector approach to the solution of the differential equations of motion of a near-Earth satellite. The method provides a good stable foundation for developing the orbital elements, thus allowing an analytic approach to be used in subsidiary algorithms. The mathematical concepts used in these algorithms are explained, and equations are developed for calculating Earth and Moon eclipses, radiation zone crossings, atmospheric density effects, solar cell decay, look angles and a geographical ephemeris. Results are presented for the IRAS satellite, and show that prediction errors of less than $\text{1}\text{1}\text{2}$ sec over one week or errors of less than 15 sec over $\text{3}\text{1}\text{2}$ months are possible.     

Developing the next phase in NASAs satellite communications program

NASAs early efforts in satellite communications development brought confidence in space technology use for improved telecommunications. New, worldwide satellite communications systems have resulted, and are now on a commercial, self-sustaining operational basis. Since 1973, NASA has conducted hundreds of user experiments and demonstrated newer technology using ATS-1, -3, -6 and CTS. Now, projections show that the commercial demand will continue to increase, soon exceeding the current technology's capacity.As a result, U.S. Space Policy affirmed in 1978 that NASA should embark again on a research and development program for satellite communications with specific, characterized goals. The resulting plan's elements include $\text{30}\text{20}\text{GHz}$ Ka-band technology, extending the current work in advanced multi-beam antennas; a narrowband system and technology study that could lead to mobile and transportable communications developments; and studies of future uses of technology in communications. The program plan and its evolution are described, followed by a report of current progress and future expectations.     

Theoretical study about the Marangoni convection in a liquid column in zero-gravity

The thermal Marangoni effect on the surface of a liquid bridge induces a convection inside the liquid. For an imposed arbitrary periodic axial circumferential temperature distribution on the liquid surface the velocity distributions in radial-, angular- and axial direction are determined theoretically by solving the linearized Navier-Stokes equations. Of particular interest is the effect of the viscosity parameter $\text{v}\text{a}{}^2$ and axial wave length to diameter ratio $\text{l}\text{a}$. It was found that the increase of viscosity decreases the magnitude of the velocity distributions and that for small axial wave length to diameter ratios the radial- and axial velocities exhibit peak values close to the free surface of the liquid. This is in a less pronounced way also true for the angular velocity, which shows for increasing moderate values $\text{l}\text{a}\text{(0 ≤}\text{l}\text{a}\text{≤ 2}$) a strong increase in magnitude and for larger axial wavelength a decrease again. For increasing axial wavelength the peak value of the radial- and axial velocity shifts towards the center of the liquid bridge, of which for a further increase a decrease of the magnitude appears.