RU2012148785A - METHOD FOR CONTROLING THE SPACE DEVICE TO THE ORBIT OF AN ARTIFICIAL SATELLITE OF THE PLANET - Google Patents

Abstract

The way to control the launch of a spacecraft into orbit of an artificial satellite of the planet, which consists in approaching the spacecraft in a given corridor for entering the planet’s atmosphere, determining in the i-th, where i = 1, 2, ..., N, time instants of the motion parameters of the spacecraft, namely : its velocity - V, the angle of inclination of the spacecraft's velocity vector to the local horizon for the trajectory of its motion - θ, the radius - of the vector - r, the density of the atmosphere - ρ, while the motion parameters of the spacecraft are determined in the area of motion in the atmosphere e of the planet, ensuring the entry of the spacecraft into the atmosphere with a roll angle γ of about 0 rad and a value of the angle of attack of the spacecraft about α, corresponding to the maximum aerodynamic quality of the spacecraft, controlling the roll angle of the spacecraft γ before departure from the atmosphere, in determining the speed of movement spacecraft in the apocenter of the transitional elliptical orbit - V and the implementation of the acceleration pulse of characteristic velocity - ∆V in the apocenter of the transitional elliptical the orbits of the spacecraft’s motion to form the orbit with the given parameters, characterized in that in the ith, where i = K, K + 1, ..., K + M, time moments at the aerodynamic drag section are predicted at the time the spacecraft leaves the planet’s atmosphere values of the speed of the spacecraft - V and the angle of inclination of the velocity vector to the local horizon for the trajectory of the spacecraft - θ when the spacecraft moves in the remaining parts of the flight in the planet’s atmosphere, respectively, with the roll angle of the spacecraft γ equal to

Claims (1)

The way to control the launch of a spacecraft into orbit of an artificial satellite of the planet, which consists in approaching the spacecraft in a given corridor for entering the planet’s atmosphere, determining in the i-th, where i = 1, 2, ..., N, time instants of the motion parameters of the spacecraft, namely : its speed - V _i, the angle of inclination of the spacecraft velocity vector to the local horizon for the trajectory of its motion - θ _i, the radius - vector - r _i, the atmospheric density - ρ _i, wherein the determination of spacecraft motion parameters is carried out on site and in motion mosfere planet providing entry of the spacecraft into the atmosphere with the value γ roll angle about 0 rad, and the value of the angle of incidence of the spacecraft about α ^*, corresponding to the maximum value of aerodynamic control of the spacecraft, the control amount of the spacecraft γ roll angle to departure from the atmosphere, in determining the speed the motion of the spacecraft in the apocenter of the transitional elliptical orbit - V _α and the implementation of the acceleration impulse of the characteristic velocity - ΔV in the apocenter of the transitional ellipse optical orbit of the spacecraft’s movement to form an orbit with specified parameters, characterized in that in the ith, where i = K, K + 1, ..., K + M, time instants in the aerodynamic drag section are predicted at the time the spacecraft leaves the atmosphere planet values of the speed of the spacecraft - V _k and the angle of inclination of the velocity vector to the local horizon for the trajectory of the spacecraft - θ _k when the spacecraft moves in the remaining sections of the flight in the planet’s atmosphere, respectively, with the value of the heeling angle of the space of the apparatus γ equal to 0 rad and γ equal to π, the heights of the apocenters of transitional elliptical orbits h _α1 and h _{α2 are} also predicted at the moment the spacecraft leaves the planet’s atmosphere when the spacecraft moves in the remaining parts of the flight in the planet’s atmosphere, respectively, with the roll angle of the spacecraft γ equal to 0 rad and γ equal to π rad, in accordance with the mathematical dependencies: $$V_k\left(\gamma =0\right)=V_i\exp \left(-\frac{{\theta }_i-{\theta }_k}{2P_{x}M-\mathrm{one}}\right)$$

, $${\theta }_k\left(\gamma =0\right)=-\sqrt{{\theta }_i^2+\frac{2}{\beta }\left(M-\frac{\mathrm{one}}{2R_x}\right)\left({\rho }_{\mathrm{one\; hundred}}-{\rho }_i\right)}$$

, $$V_k\left(\gamma =\pi \right)=V_i\exp \left(-\frac{{\theta }_i-{\theta }_k}{2P_{x}M+\mathrm{one}}\right)$$

, $${\theta }_k\left(\gamma =\pi \right)=-\sqrt{{\theta }_i^2+\frac{2}{\beta }\left(M+\frac{\mathrm{one}}{2R_x}\right)\left({\rho }_{\mathrm{one\; hundred}}-{\rho }_i\right)}$$

, $$h_{\alpha \mathrm{one}}\left(\gamma =0\right)=\frac{\mu +\sqrt{{\mu }^2-C_{\mathrm{one}}\left(\gamma =0\right)C_2\left(\gamma =0\right)}}{C_{\mathrm{one}}\left(\gamma =0\right)}-R$$

, $$h_{\alpha 2}\left(\gamma =\pi \right)=\frac{\mu +\sqrt{{\mu }^2-C_{\mathrm{one}}\left(\gamma =\pi \right)C_2\left(\gamma =\pi \right)}}{C_{\mathrm{one}}\left(\gamma =\pi \right)}-R$$

, Where $$M=\left(\frac{g{r}_i}{V_i^2}-\mathrm{one}\right)\frac{\mathrm{one}}{{\rho }_i{r}_i}$$

, $$P_x=\frac{2m}{C_x\left(\alpha \right)S}$$

, $$C_{\mathrm{one}}\left(\gamma \right)=\frac{2\mu }{r_k}-V_k^2\left(\gamma \right)$$

, $$C_2\left(\gamma \right)=V_k^2\left(\gamma \right)r_k^2{\cos }^2{\theta }_k\left(\gamma \right)$$

, V _i , θ _i , r _i and ρ _i are the current values of the spacecraft velocity, the angle of inclination of the spacecraft velocity vector to the local horizon for the trajectory of its motion, radius vector and atmospheric density, respectively; r _k is the value of the spacecraft radius vector at the spacecraft exit point from the planet’s atmosphere; R is the radius of the planet; g is the acceleration of gravity on the surface of the planet; µ is the gravitational parameter of the planet; P _x - reduced load on the frontal surface of the spacecraft; ρ ₁₀₀ - the value of the a priori density of the atmosphere at an altitude of 100 km; β is the logarithmic coefficient of change in the density of the atmosphere from height; m is the mass of the spacecraft; S is the area of the mid-section of the spacecraft; C _x - aerodynamic drag coefficient of the spacecraft, comparing the predicted value of the apocenter height h _α2 calculated at a roll angle of the spacecraft γ equal to π rad with a given height of the formed orbit h _α3 and when the condition is fulfilled: h _α2 > h _α3 increase the value of the angle of attack of the spacecraft α by a value of about Δα ', calculated by the formula: $$\Delta \alpha \text{'}\text{'}=\left({\alpha }^{**}-{\alpha }^{*}\right)\frac{h_{\alpha 2}-h_{\alpha 3}}{h_{\alpha \mathrm{one}}-h_{\alpha 3}}$$

, Where α ^* is the value of the angle of attack of the spacecraft, corresponding to the maximum value of the aerodynamic quality of the spacecraft; α ^** is the value of the angle of attack of the spacecraft, corresponding to the maximum value of the aerodynamic coefficient of drag of the spacecraft; h _α1 is the predicted height of the apocenter, calculated with a roll angle γ of the spacecraft equal to 0 rad; h _α2 is the predicted value of the apocenter height calculated at a roll angle of the spacecraft γ equal to π rad; h _α3 - given height of the formed orbit, comparing the current predicted height of the apocenter of the transitional orbit h _α2 (t _i ), calculated with a roll angle γ of the spacecraft equal to π rad with its value calculated on the previous time interval h _α2 (t _i-1 ) and if the condition: h _α2 (t _i ) <h _α2 (t _i-1 ) set the angle of heel of the spacecraft γ equal to π rad and reduce the angle of attack of the spacecraft α by about ∆α ”, calculated by the formula: $$\Delta \alpha "=\left({\alpha }_{i-\mathrm{one}}-{\alpha }^{*}\right)\frac{h_{\alpha 2}\left(t_i\right)-h_{\alpha 2}\left(t_{i-\mathrm{one}}\right)}{h_{\alpha 2}\left(t_i\right)-h_{\alpha 3}}$$

, Where α _i-1 - the value of the angle of attack of the spacecraft, calculated on the previous time interval; α ^* is the value of the angle of attack of the spacecraft, corresponding to the maximum value of the aerodynamic quality of the spacecraft; h _α2 (t _i ) is the current predicted altitude of the apocenter of the transitional orbit, calculated with a roll angle of spacecraft γ equal to π rad; h _α2 (t _i-1 ) - the predicted value of the height of the apocenter of the transitional orbit, calculated with a roll angle of the spacecraft γ equal to π rad in the previous time interval; h _α3 - set value of the height of the formed orbit, comparing the predicted value of the apocenter height h _α2 , calculated with a roll angle γ of the spacecraft equal to π rad, with a given height of the formed orbit h _α3 and when the condition is fulfilled: h _α2 = h _α3 set the value of the angle of attack of the spacecraft equal to about α ^* , which corresponds to the maximum value of the aerodynamic quality of the spacecraft, and at this set value of the angle of attack of the spacecraft is equal to about ^* spacecraft.