Method of control over spacecraft descent in atmosphere of planets

FIELD: aircraft engineering.

SUBSTANCE: invention relates to spacecraft control in atmosphere of planet by adjusting its aerodynamics. Proposed method consists in selection of conditions for changing the angle of roll to zero at changing the spacecraft from isothermal descent section (IDS) to skip path. With spacecraft in IDS, angle of roll (γ) is, first, increased to decrease aerodynamic performances and to maintain constant temperature at critical area of spacecraft surface. As flight velocity decreases angle (γ) is decreased from its maximum. In IDS, increase in aerodynamics does not cause further temperature increase over its first peak. Therefore selection of the moment of changing to γ=0 allows efficient deceleration of spacecraft at the next step of flight. The best option is the descent of spacecraft of IDS when γ reaches its maximum. Here, angle of attack is set to correspond to maximum aerodynamic performances. This increases the duration of final flight stage and deceleration efficiency. Increase in angle of attach after descent from IDS and completion of climb results in increased in drag, hence, decrease in velocity at initiation of soft landing system.

EFFECT: minimised final velocity and maximum temperature at surface critical area, lower power consumption.

2 dwg

 

The invention relates to space exploration, in particular to control the descent of a SPACECRAFT (SC) in the atmosphere of the planet, using controlled aerodynamic braking and minimizing the final speed of the spacecraft under the condition of minimum and maximum temperature in the critical region of its surface.

A known method of controlling the descent of a spacecraft in the atmosphere of the planet, using controlled aerodynamic braking and providing the lower end of the speed of the spacecraft, described in the book - Ivanov N. M., Martynov A. I. "the Motion of spacecraft in the atmospheres of the planets. M.: Nauka, 1985, pp. 168-173 - [1]. This method is to control the aerodynamic braking by changing the roll angle γ KA. The motion of the spacecraft is carried out with a constant value of angle of attack corresponding to maximum aerodynamic balancing. The method involves one-time switching of the roll angle γ with values equal to π glad to zero, which corresponds to the switching of the effective aerodynamic qualities from the minimum value (-K) to the maximum (+K).

The disadvantages of this method are as follows. Firstly, in case of using this program upravleniyaschitayut temperature in the critical region of the surface of the SPACECRAFT reach excessively high values. This is because the spacecraft with glide, to fly a mission on trajectories with multiple ricochets and, consequently, with multiple local maxima temperatures. Moreover, the absolute maximum temperatures, as a rule, coincides with the second or third local maxima. Used management program roll angle does not effectively prevent the rise of temperature after reaching his first local maximum. Secondly, the method is similar does not control the angle of attack of the spacecraft. This reduces the potential damping rate and lower maximum temperatures increase the angle of attack leads to an increase in the intensity of aerodynamic braking KA.

A known method of controlling the descent of a spacecraft in the atmosphere of the planet, using controlled aerodynamic braking and providing the lower end of the speed of the spacecraft, described in the book.M. Ivanov, A. I. Martynov, "motion Control of spacecraft in the atmosphere of Mars. Moscow, Nauka, Main editorial Board for physical and mathematical literature, 1977, pp. 159-169 - [2]. This method is the two-parameter control of the spacecraft roll angles of attack. Input kosmicheskogo the apparatus to the atmosphere is performed with the angle γ=π rad and angle of attack α, the corresponding maximum value of the balancing aerodynamic quality. On the initial flight phase switches of the roll angle γ to zero. After reaching the angle of the velocity vector to the local horizon zero value of the roll angle γ is determined from the condition for the provision of space vehicle flight on izolyatsia the plot (the plot with constant height). Then you can switch the roll angle γ to zero, providing the movement of the spacecraft on ricochetnaya trajectory with increasing altitude. On this site there is an increase in the angle of attack α from the value corresponding to the maximum aerodynamic lift coefficient KA to a value corresponding to the maximum aerodynamic drag coefficient KA.

The disadvantages of this method are as follows. First, its implementation does not control the angle of attack α of the spacecraft on the area of intensive growth temperature in the critical region of the surface. While increasing the angle of attack α in this area reduces as the velocity of the SPACECRAFT, and depend on the temperature of the heating surface. Secondly, the introduction of sovestkogo of flight does not reduce the first local maximum of the temperature, because this area starts after reaching the Tmax. Third, the descent of the SPACECRAFT with sovestkogo plot is carried out with the maximum value of the aerodynamic lift coefficient, that is, for sufficiently small values of the aerodynamic drag coefficient and the values of the aerodynamic quality that is different from the absolute maximum. In this case, there are provisions as to increase the intensity of aerodynamic drag and increase the duration of the flight. Both these factors may contribute to a lower final velocity of the SPACECRAFT.

The closest to the technical nature of the claimed method of controlling the descent of a spacecraft in the atmosphere of the planet, using controlled aerodynamic braking and providing the lower end of the speed of the spacecraft under the condition of minimum and maximum temperature in the critical region of its surface, is a method of controlling the descent of a spacecraft in the atmosphere of the planet which minimizes the maximum temperature in the critical region of the surface of the SPACECRAFT. Specified the known method described in the patent RU №2493059, publ. 20.09.2013 - [3], which is selected by the prototype. This method of controlling the descent of a spacecraft in the atmosphere of the planet lies in its space, Breakfast is the only orientation and management of aerodynamic braking, stabilization at the entrance into the atmosphere at the corners of the roll, yaw and angle of attack for maximum glide ratio, determining current coordinates and velocities of the flight of the spacecraft and the actuation means for the landing, in the process of descent continuously measure the temperature T of the external surface of the spacecraft in its critical region, for each measured temperature value T calculate the acceleration and speed of its change by calculating in time, respectively, the firstTand the secondTderivatives; upon reaching the second derivative negative valuesT<0preserving the first derivative positive values ofT>0increase the angle of attack and continue the descent to the condition of equality to zero of the first derivative ofT=0, then set the values of the Glov roll and attack, ensuring compliance with equality to zero first and second derivatives ofT=T=0at which carry out the descent of a spacecraft on the isotemperature section, upon reaching the first derivative of the negative values ofT<0establish a zero Bank angle and angle of attack for maximum value model quality, and complete the plot of inhibition of spacecraft.

The disadvantage of this method is that its implementation is not fully implemented all control in the final section of the descent of the SPACECRAFT to reduce the final speed. So, is fairly late gathering KA with isotemperature plot (ifT<0), which leads to the reduction of the length of the final segment of the flight SPACECRAFT and low intensity quenching speed. Furthermore, the method does not control the angle of attack KA immediately before the introduction of the soft landing system, which is also what leads to the decrease in the intensity of aerodynamic deceleration of the spacecraft.

The invention consists in the rational management of the roll angles of attack which minimizes the final speed of the flight SPACECRAFT while minimizing the maximum temperature in the critical region of its surface. This is achieved by introducing a new episode of the control compared to the prototype. First, the choice of appropriate conditions switching of the roll angle to zero, which ensures the vanishing KA with isotemperature site and transfer it to rikollisuus trajectory with increasing altitude. The motion of the SPACECRAFT in isotemperature site is provided by a corresponding change in angle of Bank: first there is an increase of the roll angle γ, leading to the reduction of the aerodynamic quality and maintaining a constant temperature in the critical region of the surface of the SPACECRAFT (the absence of such a control mode leads to the initial temperature decrease with further significant growth), and then decreasing the speed of the roll angle γ KA, reaching its maximum value and starts to decrease. At this stage of the flight the increase in aerodynamic quality does not lead to further increase in temperature beyond its first maximum. Therefore, when setting the zero value of the roll angle, you can achieve the most the e effective damping rate on the subsequent flight. The most rational is the gathering KA with isotemperature plot at the moment of achievement of the roll angle γ maximum value: earlier switching of the roll angle γ to a value of zero may cause further temperature rise, and a later - to reduce the duration of the subsequent section of the descent, and consequently reduce the efficiency of the damping rate of the SPACECRAFT. Secondly, at the time of descent of a SPACECRAFT with isotemperature plot set the angle of attack KA, corresponding to the maximum value of the aerodynamic quality KA (prototype - the maximum value of the aerodynamic lift coefficient KA). This is due to the need to ensure a longer final flight and more intense inhibition of CA in the atmosphere. Thirdly, the increased angle of attack KA after his departure from the isotemperature plot and complete set flight altitude increases aerodynamic drag coefficient and to a greater reduction of speed by the time of the introduction of the soft landing system.

The essence of the present invention, the method of controlling the descent of a spacecraft in the atmosphere of the planet lies in the spatial orientation of the spacecraft and the management of its aerodynamic deceleration, stabilization of the spacecraft is ri it enters the atmosphere of the planet on the roll angles γ, yaw and angle of attack α of the spacecraft, providing maximum glide ratio, the determination of the current coordinates and velocities of the flight of the spacecraft, continuous measurement of the temperature T of the external surface of the spacecraft in its critical region, we compute the velocity and acceleration of its change by computing time, respectively, the firstTand the secondTderivatives; upon reaching the second derivative negative valuesT<0preserving the first derivative positive values ofT>0increase the angle of attack α of the spacecraft and continue the descent to the condition of equality to zero of the first derivative ofT=0after which the set values of roll angles γ and attack α of the spacecraft, providing the conditions of equality to zero of the first and second proizvodi is x T=0andT=0providing the descent of a spacecraft on the isotemperature plot, while in the process of motion of a spacecraft on the isotemperature plot continuously comparing values of roll angles γ in the current tiprevious ti-1times and compute their difference Δγ by the formula

Δγ=γii-1,

where γi- the value of the roll angle of the spacecraft at time tii=1, 2, 3 ...,;

γi-1- the value of the roll angle of the spacecraft at time ti-1i=1, 2, 3 ...,

if the condition

Δγ<0,

set the value of the roll angle γ KA, equal to about 0 rad and the angle of attack α KA, corresponding to the maximum value of its aerodynamic quality, ensuring the vanishing of the spacecraft with isotemperature plot and movement ricochetnaya trajectory; carry out continuous measurement of the angle of inclination of the velocity vector θ spacecraft to the local horizon, when the condition

hi<hmax,

where hi- the current value of the altitude of the spacecraft at time tiatmosphere planet;

hmax- maximum flight altitude of the spacecraft during its movement on ricochetnaya trajectory

set the value of the angle of attack α of a spacecraft in accordance with the mathematical expression

,

where V is the current value of the speed of the spacecraft;

V0- the speed of the spacecraft when it reaches its maximum altitude at ricochetnaya trajectory after leaving the isotemperature area;

θ is the angle of the velocity vector of the spacecraft to the local horizon;

Cx0- is the aerodynamic drag coefficient at zero angle of attack of the spacecraft;

S - the area of the fuselage mid-section of the spacecraft;

M is the constant obtained after the introduction of assumptions and transform the original system of differential equations;

l, n, A - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack to the analytical mind,

to achieve the angle of attack α of the spacecraft of α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the soft landing system.

The technical result of the bretania - the method of controlling the descent of a SPACECRAFT in the atmosphere of the planet is to minimize the final velocity of the spacecraft and the minimization of the maximum temperature in the critical region of its surface, which leads to a weight reduction of thermal protection of SPACECRAFT and to decrease the required energy costs for the implementation of space missions to study the planets of the Solar system and, consequently, to increase the share of the payload in the overall weight balance of the spacecraft. The application of the proposed method, depending on the design and ballistic characteristics of SC, boundary conditions and parameters of the planets destination allows to reduce the total weight of the fuel and heat-resistant coating lander ~ 15% compared to using the prototype method.

This technical result is achieved by testing the rational management of the roll angles and attack KA, namely due to the fact that in the method of controlling the descent of a spacecraft in the atmosphere of the planet, the selected prototype consists of the spatial orientation of the spacecraft and its management aerodynamic deceleration, stabilization when entering the atmosphere of the planet on the roll angles γ, yaw and angle of attack α of the spacecraft, providing maximum glide ratio, defined the research Institute of current coordinates and velocities of the space vehicle flight continuous measurement of the temperature T of the external surface of the spacecraft in its critical region, we compute the velocity and acceleration of its change by computing time, respectively, the firstTand the secondTderivatives; upon reaching the second derivative negative valuesT<0preserving the first derivative positive values ofT>0increase the angle of attack α of the spacecraft and continue the descent to the condition of equality to zero of the first derivative ofT=0after which the set values of roll angles γ and attack α of the spacecraft, providing that the conditions for equality to zero first and second derivatives ofT=0andT= providing the descent of a spacecraft on the isotemperature area, additionally, the motion of the spacecraft on the isotemperature plot continuously comparing values of roll angles γ in the current tiprevious ti-1times and compute their difference Δγ by the formula

Δγ=γii-1,

where γi- the value of the roll angle of the spacecraft at time tii=1, 2, 3 ...,;

γi-1- the value of the roll angle of the spacecraft at time ti-1i=1, 2, 3 ...,

if the condition

Δγ<0,

set the value of the roll angle γ KA, equal to about 0 rad and the angle of attack α KA, corresponding to the maximum value of its aerodynamic quality, ensuring the vanishing of the spacecraft with isotemperature plot and movement ricochetnaya trajectory; carry out continuous measurement of the angle of inclination of the velocity vector θ spacecraft to the local horizon, when the condition is met:

hi≤hmax,

where hi- the current value of the altitude of the spacecraft at time tiin the atmosphere of the planet;

hmax- maximum flight altitude of the spacecraft during its movement on ricochetnaya trajectory

set Zn is an increase in the angle of attack α of a spacecraft in accordance with the mathematical expression

,

where V is the current value of the speed of the spacecraft;

V0- the speed of the spacecraft when it reaches its maximum altitude at ricochetnaya trajectory after leaving the isotemperature area;

θ is the angle of the velocity vector of the spacecraft to the local horizon;

Cx0- is the aerodynamic drag coefficient at zero angle of attack of the spacecraft;

S - the area of the fuselage mid-section of the spacecraft;

M is the constant obtained after the introduction of assumptions and transform the original system of differential equations;

l, n, A - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack to the analytical mind,

to achieve the angle of attack α of the spacecraft of α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the soft landing system.

The claimed method of controlling the descent of a spacecraft in the atmosphere of the planet is illustrated by the following figures.

In Fig.1 for the prototype method shows the graphs of the dependencies of the temperature T of the heating body of the SPACECRAFT in the critical region of its outer surface, karasti flying V the roll angles γ and attack α from the time of the descent into the Martian atmosphere while minimizing the maximum temperature values.

In Fig.2 for the inventive method shows the graphs of the dependencies of the temperature T of the heating body of the SPACECRAFT in the critical region of its outer surface, the velocity V and altitude h, the roll angles γ and attack α from the time of the descent into the Martian atmosphere while minimizing the final speed while minimizing the maximum temperature values.

The comparison is shown in Fig.1 and 2 are graphs we can conclude that almost the same final speed of the AC, the temperature T of the heating body of the SPACECRAFT in the critical region of its outer surface in the present invention is less than in the method prototype.

According to [2], page 194 aerodynamic coefficients of drag and lift forces with a high degree of accuracy can be approximated by the following analytical dependencies:

Cx=Cx0+Asin2(nα+l),

Cy=Cy0+Asin(nα+l)cos(nα+l).

In particular, when using forms spacecraft type bearing housing Cx0=0,2; Cy0=-0,1; A=2,3; n=1,125; l=5,625°.

For other types of forms can be used similar dependencies for other values of coefficients [2], page 194.

We will show the possibility of carrying out the invention, i.e. in the possibility of its industrial applications. A feature of the conduct of space activities in many countries is to enhance the study of the planets of the Solar system. In the framework of the Federal space program 2016-2025, provided the work to create a spacecraft to explore Mars, Venus, Jupiter, mercury, including design landers. However, one of the major problems is the development of key management technologies that reduce the mass-energy costs in all areas of interplanetary flights. In these circumstances favourable factor is the lower cost of fuel for damping of the speed during the operation of the soft landing system and weight reduction of thermal protection of SPACECRAFT. The successful solution of this problem is largely provided when placing aboard the landers control systems aerodynamic braking, using the principles of management of roll angles and attack KA set forth in the present invention.

As for the technical means to ensure control glide KA, i.e. the management of its roll angles and attack, they are known - see, for example, [1], page 37, [2], page 57, 270, and Navigation flight of the orbital complex "SALYUT-6" - "SOYUZ" - "PROGRESS"" responsible editors B. N. The PE the ditch, I. K. Bazhinov, Moscow, Nauka, 1985, Chapter 1 - [3].

Notes. 1. The applicant was placed in the Annex to the application materials rationale used them (in the description and the claims) of the mathematical expression to calculate values of angle of attack SPACECRAFT into its final section of the descent in the atmosphere of the planet that do not unnecessarily overload the description of the invention. However, if the expert deems appropriate, the applicant will not argue for its inclusion in the description.

2. According to p. 2.3.1 of the Guidelines for examination of applications for inventions from 25.07.2011 G.: the use in the claims of the sign "about" in describing the values of the numeric parameters are valid.

3. The applicant in the application materials have used two identical term "switch" the value of the roll angle γ KA (is used to describe analogues) and "set" the value of the roll angle γ KA (in the claims), as, in his opinion, is more preferable. While believing that the unity of terminology in this case is not broken.

App. Refers to the application for the invention "Method of controlling the descent of a spacecraft in the atmospheres of the planets (using controlled aerodynamic braking and minimizing the final speed of the spacecraft under the condition of minimum and maximum temperature in the critical region of the th surface (note the Applicant).

Conclusion the mathematical dependences for calculation of angle of attack on the final section of the descent of a SPACECRAFT (SC) in the atmosphere of the planet

The motion of the SPACECRAFT in the atmosphere according to the works [1, 2] is described by the system of differential equations in the velocity coordinate system with the influence of gravitational, aerodynamic, centrifugal and Coriolis forces in the assumption of the centrality of the gravitational field

Here V is the velocity of the SPACECRAFT, θ is the angle of the velocity vector to the local horizon, ε is the heading angle, r is the radius - vector connecting the center of the planet and the position of the KA, λ and φ is the longitude and latitude of subsatellite points KA, respectively, m is the SPACECRAFT mass, t is time, ρ is the air density, Cxand Cy- aerodynamic coefficients of drag and lift forces, respectively, R is the planetary radius, h is the altitude, g is the acceleration of gravity, µ is the product of the constant of gravity on the mass of the planet, S - the area of the fuselage mid-section.

The values of the control parameters α and γ can vary

0≤α≤αmax, -π≤γ≤π.

Transform the original equation (1) given the introduction of the assumptions previously used in a number of domestic and foreign operations, in particular in the works [1, 2]

h<<R, ρ=ρ0exp(-βh), Fto+FC<<Fg<<Fand,

where ρ0- p is otnesti atmosphere on the surface of Mars, β is the logarithmic rate of change of the density of the atmosphere with height, FtoFCFgFand- Coriolis, centrifugal, gravitational and aerodynamic forces, respectively.

We will consider only the final section of the descending SPACECRAFT, beginning with the time the device maximum height after the flight on ricochetnaya trajectory and ending on the date of entry into force of the soft landing system.

Specific features of the dynamics of the motion of the SPACECRAFT at this stage of the descent, due to a significant reduction in the speed of flight, are increasing gravity and reducing the aerodynamic forces acting on the SPACECRAFT. On stage after the Assembly of the apparatus with isotemperature area and maximum height of rebound gravitational forces take the values one order of magnitude larger than this model.

Using these assumptions, considering the motion of the SPACECRAFT in the plane of entry into the atmosphere and taking into account that the target area is descent with zero roll angle, convert the system of equations to a form

where M is a piecewise constant function, according to the works [1, 2].

The solution of the problem of finding the optimal SPACECRAFT control with the minimum final speed was carried out using the principle of Macs is Imola Pontryagin. Let us write the Hamiltonian H and the conjugate variables Ψi

,

,

,.

Comparing the equations for the functions H, Ψ1, Ψ3convert formulas for the adjoint variables as follows:

,.

From the transversality conditions at the end point the SPACECRAFT trajectory, it follows that

Given that the Hamiltonian does not depend explicitly on time of flight is legitimate to write the equation

H≡0.

It allows to represent the dependencies for the calculation of the adjoint variables of the form

,.

Integrating these equations, taking into account formulas (3), we obtain

,

, Ψ3(t)=a3=const.

The dependency analysis for the calculation of conjugate variables taking into account the equality to zero of the Hamiltonian has shown that Ψ1(t) is negative monotonically increasing function, reaching the end point of the trajectory values of -1; Ψ2(t) is a positive monotone decreasing function, reaching the end point of the trajectory values of zero; Ψ3(t) is a continuous function is th, having a negative value.

From the condition of maximum of the Hamiltonian determine the law of variation of the aerodynamic drag coefficient Cxat the final stage of the flight SPACECRAFT

Cx=-signΨ1.

Given that the associated variable Ψ1is negative, the aerodynamic coefficient Cxwill take the maximum value at the end of the flight path.

To determine the dynamics of changes in the value of Cxdivide the first equation of system (2) on the second and get the following equation

.

As a result of its decision will write down the formula for determining the aerodynamic coefficient Cxdepending on the changing values of V and θ:

where V0- the speed of the spacecraft when it reaches its maximum altitude at ricochetnaya trajectory after leaving the isotemperature plot.

According to the works [1, 2], the aerodynamic coefficients of drag and lift forces with a high degree of accuracy can be approximated by the following analytical dependencies:

Cx=Cx0+Asin2(nα+l),

Cy=Cy0+Asin(nα+l)cos(nα+l).

For landers-type bearing housing Cx0=0,2; Cy0=-0,1; A=2,3; n=1,125; l=5,625°.

Considering elisavietta formula (4) is converted to the following:

,

The analysis of this equation showed that the angle of attack α in the area of flight KA increases monotonically with increasing intensity and reaches the end of the segment size equal to ≈70÷85°, which corresponds to the maximum value of the aerodynamic coefficient of drag.

Sources of information

1. Ivanov N. M., Martynov A. I. "the Motion of spacecraft in the atmospheres of the planets. M.: Nauka, 1985, pp. 168-173.

2. N. M. Ivanov, A. I. Martynov, "motion Control of spacecraft in the atmosphere of Mars. Moscow, Nauka, Main editorial Board for physical and mathematical literature, 1977, pp. 159-169.

The method of controlling the descent of a spacecraft in the atmosphere of the planet, lies in its spatial orientation and management of aerodynamic deceleration, stabilization of the SPACECRAFT (SC) when entering the atmosphere of the planet on the roll angles γ, yaw and angle of attack α for maximum glide ratio in the determination of the current coordinates and velocities of the flight SPACECRAFT, continuous measurement of the temperature T of the external surface of the SPACECRAFT in its critical region, we compute the velocity and acceleration of the temperature changes by computing time, respectively, the firstand the seconda derivative is, however, achievement of the second derivative negative valueswith retention of the first derivative positive valuesincrease the angle of attack α KA and continue the descent to the condition of equality to zero of the first derivativeafter which the set values of roll angles γ and attack α KA, ensuring compliance with equality to zero first and second derivativesandproviding the descent KA on isotemperature section, characterized in that during the motion of the SPACECRAFT on the isotemperature plot continuously comparing values of roll angles γ in the current tiprevious ti-1times and compute their difference Δγ according to the formula
Δγ=γii-1,
where γi- the value of the roll angle CA at time ti, i=1, 2, 3 ...,
γi-1- the value of the roll angle CA at time ti-1, i=1, 2, 3 ..., and if the condition is true
Δγ<0
set the value of the roll angle KA γ is equal to about 0 rad and angle of attack KA α corresponding to the maximum value of its aerodynamic qualities, providing a gathering KA with isotemperature plot and movement ricochetnaya trajectory, carry out continuous measurement of the angle θ of inclination of the vector soon the ti KA to the local horizon, and if the condition is true
hi< hmax,
where hi- the current value of the altitude of the SPACECRAFT at time tiin the atmosphere of the planet,
hmax- maximum altitude SPACECRAFT during its movement on ricochetnaya path,
set the value of the angle of attack α in accordance with the mathematical expression
,
where V is the current speed value KA,
V0- the speed of the SPACECRAFT when it reaches its maximum altitude at ricochetnaya trajectory after leaving the isotemperature plot
θ is the angle of the velocity vector of the SPACECRAFT to the local horizon,
Cx0- is the aerodynamic drag coefficient at zero angle of attack KA,
S - the area of the fuselage mid-section KA,
M is the constant obtained after the introduction of assumptions and transform the original system of differential equations;
l, n, A - constant coefficients of the approximation of the dependency of the aerodynamic coefficients on the angle of attack to the analytical mind,
and to achieve the angle of attack KA α α*, corresponding to the maximum value of the aerodynamic coefficient of drag, fly a mission with this value of angle of attack α* before commissioning the soft landing system.



 

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13 cl, 16 dwg

FIELD: chemistry.

SUBSTANCE: invention is referred to means of matching and separation of pneumohydraulic systems of space objects (SO). The detachable device has forwarding studs (1) with fixing elements and division gears. Studs may be fixed on a detached part (9) of the SO, and fixing elements - on the core vehicle (11) of the SO, and division gears (engines) - on the launch vehicle. Transit line is equipped with the first (4) and the second (5) adapters which form a detachable joint in the shape of adapter (4) top (6) and adapter (5) flange (7). The top (6) enters the flange (7) core space, and their joint is sealed with inserts (8). The adapter (5) is attached to the detached part (9) of the SO with the help of a flange (7) and fixturing components (10), and the adapter (4) is fixed on the core vehicle (11) of the SO with a pylon (12). Supporting part (13) of the pylon is fixed on the core vehicle (11) of the space object with the help of elements (10), and an adjustable part (14) - on the screwed shank (15) of the adapter (4) with the screw nuts (16). Oval holes (17) provide regulation of relative position of the parts (12) and (14) on installation sites of fixturing elements (10). Displacement of the pylon (12) in relation to the adapter (4) is fixed with two screw nuts (16) by means of their screwing along the tail (15). Studs (1) cylinder part (18) is longer than disconnection stroke of adapters (4) and (5). Thus, deformation by the disconnection is not possible. After SO parts connection, the adapter (5) is fixed on the detached part (9). Then the first adapter (4) is attached to it and it the pylon (12) is installed on it. The pylon (12) is fixed on the core vehicle (11). Then transit pipelines are attached to the adapters and adapters' conjunctions are technologically unfixed.

EFFECT: increased safety of the detached device by minimal mass expenses.

2 cl, 3 dwg

FIELD: space and aeronautics.

SUBSTANCE: invention is referred to matching missions, in particular that of a manned space vehicle from an international space station. The mode involves active space object (ASO) final orbit insertion by a launch vehicle with deviations along the longitude of ascending node, and passive space object (PSO) orbit inclination. Thereby, ASO and PSO latitude argument orbits mismatch is prescribed. In orbits intersection in a direction orthogonal to plane, leading outs do maneuvers for the elimination of the stated deviations. Thereby, they form a speed pulse, which leads to the match of orbit planes of ASO and PSO.

EFFECT: reduced amount of time for matching of ASO and PSO without additional corrections of the orbit of PSO.

9 dwg, 1 dwg, 2 tbl

FIELD: transport.

SUBSTANCE: invention relates to docking jobs, in particular, to docking of spaceship with international space station. Proposed method comprises using carrier rocket to place active spaceship (ASS) into target orbit with orbit ascending section altitude deviation but with preset latitude departure argument mismatch. Then, first maneuver of ASS is carried out in direction perpendicular to placing plane to change orbit ascending section altitude. Further, second maneuver is performed in opposite direction to eliminate mismatch in target orbit inclination originated after said first maneuver. Said maneuver is performed with ASS engine thrust throttling to create additional orbit ascending section altitude deviation and align orbital planes of objects to be docked. Formation of required initial angular mismatch between said objects, period approach of said objected decrease up to one pass.

EFFECT: decreased docking interval without additional orbit corrections.

7 dwg, 2 tbl

FIELD: physics; control.

SUBSTANCE: invention relates to controlling movement of a spacecraft fitted with a heat radiator and a solar panel. The method includes flying the spacecraft on an orbit around a planet and turning the solar panel in a position corresponding to the alignment of the normal to the working surface of the solar panel with the direction towards the Sun; performing orbital orientation of the spacecraft, where the plane of rotation of the solar panel is parallel to the plane of the orbit of the spacecraft and the solar panel is located relative to the plane of the orbit on the side of the Sun; determining the maximum value of the angle between the velocity vector of the spacecraft and the perpendicular to the transverse axis of rotation of the solar panel, passing through the surface of the radiator; determining the orbital altitude of the spacecraft and the angle between the direction towards the Sun and the plane of the orbit of the spacecraft; based on the orbital altitude and the angle, determining the orbit passes where the duration of the illuminated part of the pass exceeds the difference between the orbiting period of the spacecraft and the required duration of the heat release by the radiator on the pass; on the said orbit passes, when the spacecraft passes through the illuminated part of the pass, the solar panel is turned around the transverse axis of rotation until the intersection of the line passing through the region of the surface of the radiator facing the Sun and directed towards the Sun with the solar panel; turning the solar panel around the longitudinal axis of rotation until the angle between the normal to the working surface of the solar panel and the direction towards the Sun assumes a minimum value. The said solar panel rotations are performed within a calculated time interval.

EFFECT: high efficiency of the radiator by creating conditions for natural cooling thereof during eclipse of the solar panel for any altitude of an almost circular orbit of the spacecraft.

5 dwg

FIELD: transport.

SUBSTANCE: invention relates to space cable systems (SCS) and can be used for the transfer of SCS to a spinning mode in the orbit plane without the application of jet engines. SCS development is executed from its initial compact state in the circular orbit by the repulsion of objects at a low relative speed. SCS end weights are connected by a cable, its length being varied by a cable feed-haul-in device arranged on one of the end objects. The objects are separated by a vector of local peripheral speed, for example, by a pusher. The objects are driven by a start pulse to separate the objects in practically free paths at the free feed of the cable. The cable development is terminated by the SCS transfer to a stable mode of associated pendulum motion at the stretched preset-length cable. At a definite range of angular phases of this mode the SCS objects are stretched by hauling in the cable at a definite constant speed. This results in changing the SCS into the spinning mode at a preset power integral and fixed final end of the cable.

EFFECT: relaxed weight-size constrictions of SCS, enhanced performances.

8 dwg

FIELD: physics, navigation.

SUBSTANCE: group of the inventions relates to control of angular motion of space vehicle (SV). The method includes additional generation of signals for assessment of orientation angle and angular velocity of rotation of space vehicle. Also the reference signals of the orientation angle, angular velocity and control assessment signal are generated. For the named orientation angle and angular velocity their differences with their assessed signals, and also the difference with their reference values are determined. The difference of control signal and its assessed value and, at last, the signal of correction of the signal of assignment of mathematical model and the signal of assessment of external noises using the respective formulas are determined. On this base the signals of assessment of orientation angle and angular velocity of space vehicle are determined, which are used for space vehicle control. The device in addition contains the reference model of the basic circuit of orientation of space vehicle and other necessary devices and connections.

EFFECT: improvement of orientation accuracy and operational reliability in case of failures of orientation angle sensor and sensor of angular velocity of space vehicle rotation.

2 cl, 2 dwg

FIELD: physics.

SUBSTANCE: invention relates to controlling orientation of an artificial earth satellite with solar panels. The disclosed method includes performing necessary turning of the artificial earth satellite along with solar panels and, separately, the solar panels about a first and a second axis. The antenna of the artificial earth satellite is directed towards the earth and the normal to the solar panels is directed towards the sun. Independent programmed turns about the first and second axes of the artificial earth satellite are performed in intervals of uncertainty of orientation of the artificial earth satellite on shadow orbits. In different versions of said turns, after the first turn, the artificial earth satellite is held in an intermediate position and normal orientation of the artificial earth satellite is then restored. This improves the accuracy of predicting movement of the artificial earth satellite on shadow orbits and accuracy of measuring the range to the artificial earth satellite.

EFFECT: high accuracy of determining navigation-time data on navigation artificial earth satellites by consumers.

4 cl, 12 dwg

FIELD: aircraft engineering.

SUBSTANCE: invention relates to in-flight control over spacecraft equipped with heat radiator and solar battery. Proposed process comprises spacecraft flight in orbit around the planet with solar battery turn to position corresponding to normal to solar battery working surface directed to the Sun. Spacecraft orbital orientation is constructed whereat solar battery spinning plate is parallel with spacecraft orbit plane while solar battery is located on the Sun side relative to orbit plane. Spacecraft orbit altitude and angle between direction to the Sun and spacecraft orbit plane are defined. Magnitude of said angle (β*) is defined whereat duration of turn shadow section equals the necessary time of radiator heat release in said turn. Orbit turns are defined wherein current magnitude of said angle is larger than β*. In said turns, solar battery is turned around crosswise and lengthwise rotation axes unless shadowing of solar battery radiator. Note here that minimum departure of orientation of solar battery working surface to the Sun. Spacecraft orbital flight is conducted in near-circle orbit at altitude not exceeding a definite design value.

EFFECT: higher efficiency of radiator with solar battery shadowed at whatever position of spacecraft on orbit turn.

3 dwg

FIELD: transport.

SUBSTANCE: invention relates to control over spacecraft, particularly, to holding of geosynchronous spacecraft in preset are of stay and collocation with the other geostationary spacecraft. Proposed method comprises determination and correction of initial inclinations and longitude of injection orbit ascending node with allowance for epoch of spacecraft placing in orbit and term of its active existence. Note here that the time of beginning of operation in geostationary orbit when spacecraft orbit inclination reaches maximum permissible value iper. area. The latter corresponds to permissible reach in latitude at the boundary of nominal spacecraft stay area in altitude. Stable and minimum eccentricity magnitudes are defined. Eccentricity vector is corrected so that it equals the nominal value for spacecraft collocation and spacecraft orbit apse line is aligned with that of nodes. Spacecraft active collocation is executed at changing the inclination from 0 to iper without interaction with adjacent spacecraft control centres. At inclination larger than iper , eccentricity is increased to minimum with setting of Laplace vector in direction from the Sun. Note here that eccentricity vector is not corrected unless the end of spacecraft active existence term termination. At inclinations larger than iper, eccentricity vector equals modulo and is spaced apart relative to eccentricity vectors of the other spacecraft.

EFFECT: decreased power consumption for stay area and collocation of geostationary spacecraft.

9 dwg

FIELD: instrumentation.

SUBSTANCE: invention relates to control of movement of space vehicle (SV). According to the proposed method, thrusts of correction engines (CE) (control accelerations) are determined as per total changes in a period of SV revolution from correction to correction. The latter are performed with one and the same CE and a thrust level of those CE is estimated. For reliable understanding of thrusts of a pair of mutually opposite located CE, series control activations of the same pair are performed from time to time with equal pulses. Discrepancy as to total pulse of thrust is entered in equal parts with an opposite sign to the implemented pulses. As a result, reliable thrust levels of CE in operation are obtained.

EFFECT: reduction of costs and improvement of CE thrust determination accuracy as per the data of trajectory measurements, as well as improvement of SV orbit correction accuracy.

FIELD: transport.

SUBSTANCE: invention relates to spacecraft (SC) motion control using solar radiation pressure forces distributed over SC working zones. The latter are formed as flat parallel optically transparent droplet flows. Distance between droplets of R radius in each flow in its lengthwise direction (Sx) and frontal-lateral direction (Sy) is divisible by 2R. Number of flows is n=(Sx/2R)1. By mutual bias of flows in direction of their motion for 2R distance droplet mist flows are generated in number of m=(Sy/2R)1. Each of the mentioned flows is biased relative to previous flow for 2R distance in frontal-lateral direction. Thus opacity in frontal-lateral direction and transparency in direction of plane perpendicular to a flow is created. Unit distributed light pressure force is regulated by changing radius and number of droplets coming to point of it application in unit time. Total action value is regulated by changing number of droplet jets.

EFFECT: increase of efficiency of light pressure distributed external forces usage by means of decreasing their disturbing effect on relative SC motion.

3 dwg, 1 tbl

FIELD: chemistry.

SUBSTANCE: invention can be used to move space debris from working orbits to recycling orbits. The method includes taking towing spacecraft and a self-contained docking module into the region of an orbit from which space debris is to be cleared. The sequence of removing space debris is selected by comparing a criterion, for example the probability of the space debris colliding with other space objects, for each space debris. Compensation for accumulated errors of motion parameters of the towing spacecraft during previous manoeuvres, as well as the pointing system is distributed between correcting pulses of the towing spacecraft at the long-range guidance step and of the self-contained docking module at the self-guidance portion.

EFFECT: high efficiency of removing space debris from working orbits.

FIELD: engines and pumps.

SUBSTANCE: proposed method comprises generation of data signal in every channel corresponding to a definite combustion chamber as the difference between command signal and feedback signal, generation of control system channel cut-off signal as the signal of setting actuator rod on said channel to mid position. Channel cut-off signal is generated in case the data signal modulus integral calculated at preset-duration time interval exceeds the preset threshold. Note here that command signals of other channels are generated as sums of or difference in pitch, yaw and bank control signals and cut-off channel feedback signal with coefficients depending on cut-off channel number so that required summed pitch, yaw and bank control moments are actuated.

EFFECT: higher reliability at fault of control system channel.

4 dwg

FIELD: space engineering; designing spacecraft motion control systems.

SUBSTANCE: proposed method is performed by information of orientation unit to Sun by introducing the orbit parameters into on-board computer followed by calculating the Sun position in observation field of orientation unit for each point of orbit for orientation of axes in orbital coordinate system; search angular velocity is set for spacecraft to ensure capture of Sun by observation field of orientation unit, after which angular velocity is decreased to zero ensuring position of Sun in observation field of orientation unit. Then spacecraft is turned in such way that Sun should move to required initial point; turning the spacecraft to preset points is continued for each orbital point.

EFFECT: reduced mass; simplified construction of spacecraft due to reduced number of instruments and units; extended field of application.

3 dwg

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