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.     

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.     

On the limiting mass of a dense stellar matter

In the present work, we have derived an expression $\text{M}{}_{\mathrm{L}}\text{? 4.738 M}{}_{\mathrm{\odot }}$ (M_⊙ = 1.985 × 10³³ g = mass of the Sun) giving the “limiting” value of the mass of a dense stellar matter, by introducing the concept of nuclear size correction in the theory of relativistic Thomas Fermi model for a compressed atom. We find that $\text{M}{}_{\mathrm{L}}\text{? 5.1571 M}{}_{\mathrm{<mtext>Ch</mtext>}}\text{and}\text{=3.2750(M}{}_{\mathrm{<mtext>O</mtext>}}\text{)}{}_{\mathrm{<mtext>Prev</mtext>}}$ [M_Ch and (M_O)_Prev denote respectively the Chandrasekhar and author's “limiting” masses]. By making a comparative study with those of previous results it has been shown that our present treatment would provide satisfactory results for the density ranges from ? ? 10⁸ up to $\text{? ? 10}{}^{11}\text{g}\text{/}\text{cm}{}^3$. Other results of cognate interest in the non-relativistic regime 10³ < ? ? 10⁵ (without the nuclear size effect) are presented. The astrophysical implications of the results are mentioned.     

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.     

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.     

Analyse du fonctionnement d'un propulseur a éjection modulée

ONERA developed, for studying the response of a propellant to a pressure or velocity fluctuation, an experimental rocket engine whose nozzle throat area can be modulated by a toothed disk.The paper presents a linearized theory of the functioning of this engine in the low frequency domain, i.e. when there is no wave propagation within the combuster.To describe the functioning of this motor, the Ryazantsev-Novozhilov method, which assumes that the gas response is instantaneous, is used. This analysis takes into account the erosion and radiation effects, the combustion efficiency and the thermal losses through the walls.Two particular cases are described, for two values of the Damköhler parameter $\text{D}{}_1\text{=}\text{t}{}_{\mathrm{c}}\text{t}{}_{\mathrm{th}}$, where t_c is the residence time in the combuster and t_th the characteristic thermal time of the heat penetration into the solid propellant. These two cases correspond, one to a classical propellant D₁ > 1, the other to a particular propellant of low burning rate ($\text{J}{}_{\mathrm{b}}\text{? 0.2}\text{to}\text{0.4}\text{mm s}{}^{\mathrm{?1}}$) D₁ < 1. The stability conditions are analysed as well as the pressure amplitute and phase as a function of the nozzle throat modulation frequency.Still in linearized theory, the complete solutions of the problem are presented, using a method of numerical resolution.     

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.     

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.     

On a lunar space elevator

We consider a space elevator system for lunar surface access that consists of a space station in circumlunar orbit, a cable reaching down to some meters above the surface and a magnetically levitated vehicle driven by a linear motor. It accelerates the load to be lifted to the speed of the cable end. Loads to be delivered are either put on the vehicle and slowed down by it or they are slowed down by a sand braking technique in a mare terrain. It is technically possible to operate this transport system nearly without fuel supply from Earth. We calculate various steel cable dimensions for a static stress maximum of $\text{1}\text{5}\text{th}$ of the tensile strength. The process of takeover is considered in detail. Five ways of eliminating the adverse large cable elongation due to the load are described. The touchdown process and behaviour of the cable after disconnection are analysed. The positive difference between the speed of the load at takeover and cable end can excite a large inplane swing motion. We propose to damp it by a dissipative pulley that hangs in a loop of wire leading to the ends of two beams mounted on the space station tangentially to the orbit, the pulley's core being connected with the load. Roll librations are damped by energy losses in the elastic beams; damping can be reinforced by viscous beam elements and/or controlled out-of-plane motions of the beams. We argue in favour of the possibility of fast deployment. The problems of vehicle vibrations and agglutination at sand braking blades are underlined and their combined experimental investigation is suggested.     

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.