Way down the accelerator of a space rocket in the landing zone and device for its implementation

 

(57) Abstract:

The invention relates to rocket and space technology and can be used to bring the exhaust accelerator first-stage space rocket "proton-M" in a limited area of impact to reduce the effect of ILV on the ecological state of the area of operation. The invention solves the problem of increasing management efficiency accelerator according to his casting in the area of the fall, reducing loads on the governing bodies and the area up to a radius of not more than 100 caliber rockets. Way down the accelerator of a space rocket /LV/ landing zone includes the stages of stabilization, targeting and controlled descent using aerodynamic forces. When the accelerator of the first stage is separated together with the wall of the caudal compartment of the second stage on the stage of stabilization, while achieving high-speed pressure 40 kg/m slow down and stabilize the accelerator stage in the engine forward position by means of symmetrical relative to the longitudinal axis ILV pairwise synchronous rotation prednisalone aerodynamic control surfaces of the tail compartment of the second stage. On stage, aiming at the area of incidence of the accelerator include the beacon, and onboard computational point of incidence and the location of the beacon, and with elektrogidrodinamicheskie actuators control the position of the aerodynamic rudders for yaw and pitch depending on the magnitude of the error. The device release the accelerator ILV in the landing zone contains aerodynamic rudders and elektrogidrodinamicheskie drives, with aerodynamic control surfaces made telephonetype, mounted on made in one piece with the first accelerator stage and located offset relative to the lateral sides of the first stage on poluga step panels of the caudal compartment of the second stage and kinematically linked with actions synchronously in pairs symmetrically to the longitudinal axis of ILV and elektrogidrodinamicheskie actuators associated with the radar, with the possibility of action in proportion to the magnitude of the error of the estimated point of incidence and the position of the beacon. 2 S. p. f-crystals, 8 ill.

The invention relates to rocket and space technology and can be used to bring the exhaust accelerator first-stage space rocket "proton-M" in a limited area of impact to reduce the effect of ILV on the ecological state of the area of operation.

Known methods of control using this model with the additional authorities [1]

Known, taken as a prototype, the method of descent of the aircraft for space purposes in the landing zone, comprising the steps of stabilization, aiming and managed using aerodynamic forces of descent to a landing area used on the spacecraft "Vostok", "Voskhod", "Union", "Bor", "Spiral", and others [2]

Known variant of the aerodynamic control surfaces on the principle of brake spoilers [3] i.e., using aerodynamic forces acting parallel to the longitudinal axis of the accelerator.

This method of controlling the descent requires the installation of a powerful hydroelectromechanical drives to rotate and hold the brake planes. The main disadvantage is that the shoulder aerodynamic forces relative to the center of gravity (CG) is not more than 2.5 m ie 5.6 times less than in the case of control of aerodynamic forces in the transverse plane. These negative aspects of such Executive bodies cause the low efficiency of their use on LV.

Also known variant management working accelerator aerodynamic forces in the transverse plane, which is implemented in the device PH Saturn (USA) [ part of the first accelerator stage.

There is a small shoulder of the effects of aerodynamic forces relative to the CG of the dry accelerator significantly (almost six times) affects the process of stabilization of the accelerator with his entry into the atmosphere and for this reason, the controllability of the accelerator at high speed pressure is not effective, since a height of about 25 km. Increase as the square of the aerodynamic control surfaces increases the weight of accelerator (more than 6 times) and operational characteristics of the missile.

The aim of the invention is to improve the accuracy of the landing rockets on a limited area (to reduce the size of areas falling accelerator ILV with a flexible management system) with a minimum weight cost.

This is achieved by the method of descent accelerator space rocket in the planting area using aerodynamic forces, including the stages of stabilization, targeting and controlled descent, separate the accelerator first stage together with the wall of the caudal compartment of the second stage, on the stage of stabilization, while achieving high-speed heads 48 kg/m2(at an altitude of about 36 km) slow down and stabilize the accelerator first stage engine forward symmetrical relative to the longitudinal axis ILV pairs of sinkronisasi in the area of the fall of the accelerator include the beacon, and then use the on-Board computer determines the skew of the side and the distance between the calculated point of incidence and the location of the beacon and using elektrogidrodinamicheskie actuators control the position of the stabilizers according to the yaw and pitch depending on the magnitude of the error.

This goal is achieved by the fact that the device release the accelerator ILV in a landing zone containing aerodynamic rudders and elektrogidrodinamicheskie drives, aerodynamic handlebars made telephonetype mounted on the panels of the caudal compartment of the second stage, arranged offset to the side blocks in the first stage of polyglossia and made in one piece with the first accelerator stage and kinematically associated with actions synchronously in pairs symmetrically to the longitudinal axis of ILV and elektrogidrodinamicheskie actuators associated with the onboard radar or homing head with the possibility of action in proportion to the magnitude of the error of the estimated point of incidence and the position of the beacon.

The proposed method is illustrated by the diagram of Fig flight. 1.

Way down the accelerator ILV in the planting area using this model silgo descent.

Features of the proposed method lies in the fact that to increase the accuracy of landing on a limited area, after the end of the active site 1 launch accelerator first stage carry out its unit 2 together with the tail compartment of the second stage. When entering the atmosphere, on the stage of stabilization (at an altitude of about 36 km) in area 3 achieve high-speed pressure of 40 kg/m2slow down and stabilize the accelerator 4 first-stage engine forward symmetrical relative to the longitudinal axis ILV pairwise synchronous rotation of the aerodynamic control surfaces, tail compartment of the second stage, acting along the axis of the accelerator.

At the stage of targeting 5 in the zone of incidence 6 of the accelerator include beacon 7, and then using a computer to determine (calculate) the error on the side and the distance between the calculated point falls 8 and the location of the beacon 7 and on the site of the fall 9 using elektrogidrodinamicheskie actuators control the position of the aerodynamic rudders for yaw and pitch depending on the magnitude of the error, the aerodynamic forces in the transverse plane, on the understanding that it is necessary to choose the limiting error of the dispersion of the first accelerator stage when padanian fall 6 without the use of the proposed method and the area falling 10 under the proposed method.

In Fig. 2-4 depicts the device, General view; Fig. 5 nodes and a fragment of a block diagram of the SU.

The tail compartment 11 (Fig. 2) do the second step is performed together with the spent catalyst of the first stage 12 as a whole. Panel 13 of the tail compartment 11 do the second stage zakomponovany with the Executive bodies of the six aerodynamic control surface 14. They are located in the front of the engine compartment of the second stage, which is shared with his step on guide and taken together with the accelerator of the first stage. Aerodynamic handlebars 14 are relatively outboard fuel tanks "G" 15 the first stage offset 30about. When deriving the orbital cargo plane aerodynamic rudders are located downstream (position 16).

After about 30 to 40 seconds after separation of the spent accelerator first stage from the second plane rudders are installed in extremely rotated position (60about) position 17, with three plane aerodynamic rudders are rotated in one direction and three plane in the opposite, keeping symmetrically to the longitudinal axis of the accelerator. In this position they are at the entrance to the upper layers of the atmosphere are as brake flaps, promote fast stabiliz km aerodynamic handlebars 14 are installed along the stream 16 and in accordance with the received information (mapping beacon 7, installed in the area of the fall of 6, and the analysis of the calculated errors of the landing area 10 from 19 location beacon). By changing the angle 20 installation of the control changes the angle of attack of the respective wings, it creates an aerodynamic force control on the shoulder 21 of the CT 22 is not less than half the length of the stage (fifteen meters for "proton").

Each plane aerodynamic wheel 14 has an area of about 7% of the area of the midsection (1.5 square meters for "proton") and is mounted rotatably around the axis 23 at an angle (20) of plus or minus 60 relative to the longitudinal axis 24 of the accelerator. The axis of rotation 25 is within 50% of the chord of the tail, which has the correct form of a trapezoid with a wingspan of about 20% of the diameter of the midsection (1.5 m for "proton").

In Fig. 2 shows the basic geometrical relations detachable accelerator first stage motor compartment of the second stage that has aerodynamic handlebars.

The axis of rotation 25 of the steering wheel, which is a continuation of the side member 26, mounted on two bearings 27 and 28 by bearings 29 and 30 of the needle type. Efforts on piers 27 and 29 acting on the aerodynamic forces on the control surfaces 14, perceived root 28 and support at support 27. The force on the root of the PBO is the Torah stage.

While increasing the casing 33, the stringers and formers of this zone compartment, especially in the field installation of the power boxes 31 and bearing units for extreme support 27 of the steering wheel 14.

Using electrothermomechanical actuator 34, which is connected with a rocker 35, fixed on the axis of plumage, is an Autonomous rotation of each aerodynamic steering signals SU.

Each actuator 34 provides a force on the rocker 35 steering is supposed to be 8 so All drives (6 PCs) operate autonomously from battery 36 pressure balloon type (from 3 to 6 pieces), electrically connected between an initial pressure of approximately 200.250 ATM.

To control higher system 37 of the actuator, using the electrical signals necessary energy is provided by existing BHB 38 aboard the first accelerator stage.

Management conversion accelerator first stage to the limited size of the drop is performed in the following way.

The angle of attack (Fig. 2) accelerator first stage at the time of separation steps is approximately equal to 0about, while the stable equilibrium is achieved when the angle of attack of 180about. From any disturbances after separation, the accelerator 12 starts to rotate in m 65aboutC.

This rotation lasts about 160 C until the moment of entry into the dense layers of the atmosphere (q about 600 kg/m2N about 30 km), where due to the static stability of rotating enters fluctuations relative to the balancing position. The control of the rudders 14 in the mode damping is fast enough to suppress these oscillations, i.e., to provide small angles between the longitudinal axis and the velocity vector. According to preliminary data source, the efficiency of the rudder 14 is sufficient for the output of the accelerator on the angle of attack in the range of 180 plus or minus 30about. Trajectory calculations show that at a steady controlled flight with a height of about 36 km and with the true angle of attack of 180 plus or minus 25aboutthe drop point may be rejected by the value of (Delta) plus or minus 4 km. This determines the limit maneuvering abilities on a downward trajectory.

The operation of the guidance system.

To provide controlled flight must (in addition to the angular stabilization) to determine aboard the coordinates of the accelerator relative to the sighting point of the fall 10. The use of this traditional inertial hydraulic excluded due to unacceptably large angles and angular skoracki (expensive) instrumentation and on-Board computer.

The proposed use of perspective in weight and cost of the radar system consisting of stationary passive homing head or 40 radar onboard the accelerator and active beacon 7 in the landing zone with a hemispherical radiation pattern, which is applied in the proposed method the device. In the homing head (radar) 40 is electronically scanned in two orthogonal planes to determine the angles between the longitudinal axis 24 of the degree and direction 41 on the beacon. The initial accuracy of the estimates of these angles should be about 0,5about.

In the channel of the roll should be provided only damping, i.e. Pets slow the spread of the accelerator angle roll in the process guidance.

The device has the following instrument components (excluding reserve):

for the system angular stabilization 42:

three angular velocity sensor 43 (DOS type BDG-36) one in each channel,

it may take two accelerometer 44 to stabilize pitch and yaw, the need for which is determined later in the analysis of sophisticated aerodynamic characteristics, taking into account tolerances;

for system induced the napravlennosti and apparatus 46 signal processing and generation of the error in the direction of motion. For both systems require a simple on-Board computing device 47. The total weight of additional equipment about 30 to 40 kg.

Thermal management system rocket the proposed method will reduce the variation in the trajectory of the passive flight spent the first accelerator stage at the heights of his entrance into the dense layers of the atmosphere (N approximately 36 km).

In this regard, it can be expected that on the flight from N about 35 km to the ground (H= 0), taking into account the impact on the accelerator external factors, aerodynamic control surfaces 14 with the adopted geometrical ratios, reject the results of the calculation control signals on-Board accelerator using the measured radar mismatches can bring the accelerator on a limited area of the drop radius variation (Delta) is not more than 0.6 km

1. Way down the accelerator of a space rocket in the landing zone, comprising the steps of stabilization, targeting and controlled descent using aerodynamic forces, characterized in that the separated accelerator first stage together with the wall of the caudal compartment of the second stage, on the stage of stabilization when reaching the velocity head is correctly longitudinal axis of the rocket pairwise synchronous rotation prednisalone aerodynamic control surfaces of the tail compartment of the second stage, at the stage of targeting include on-Board radar and using on-Board digital computing machine determine the skew of the side and the distance between the calculated point of incidence and the location of the beacon and using elektrogidrodinamicheskie actuators control the position of the aerodynamic rudders for yaw and pitch depending on the magnitude of the error.

2. Device for draining off the accelerator missiles in a landing zone containing aerodynamic rudders and elektrogidrodinamicheskie actuators, characterized in that the aerodynamic handlebars made telephonetype mounted on the panels of the caudal compartment of the second stage, arranged offset to the side blocks of the first stage on the floor-corner-step and made in one piece with the first accelerator stage, and kinematically associated with actions synchronously in pairs symmetrically to the longitudinal axis of the rocket, and elektrogidrodinamicheskie actuators associated with the radar with the possibility of action in proportion to the magnitude of the error of the estimated point of incidence and the position of the beacon.

 

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Dirigible // 2385818

FIELD: transport.

SUBSTANCE: invention refers to air navigation equipment. Dirigible includes a frame with the covering forming a cavity, ballonets with lighter-than-air gas, which are attached to the frame, stabiliser, propeller, crew compartment, passenger compartment and mooring device. In lower part of cavity there horizontally located are at least two rotors rotating in opposite directions. Rotors can be installed across the axis and lie in one and the same plane. Also rotors can be installed along the axis or one above the other.

EFFECT: increase of flight stability.

4 cl, 5 dwg

FIELD: transport.

SUBSTANCE: invention relates to the field of aviation, more specifically to a method for balancing fuel reserve in aircraft wing tanks during ground operation. The method consists in the following: balancing is executed via fuel pumping from the wing tank which contains more fuel to the central tank of an aircraft. In this process, to compensate a difference in fuel reserve between wing tanks the aircraft pump for fuel supply from left wing tank is switched on, centralised dump valve and central tank filling valve is opened, and fuel is fed under pressure to the central tank through central tank filling valve. Minor fuel part is returned from the central tank to the left wing tank by fuel jet pump.

EFFECT: prevention of transverse imbalance between left and right wing tanks.

4 cl, 1 dwg

FIELD: transport.

SUBSTANCE: invention relates to aircraft engineering, particularly, to method of equalising fuel capacity in aircraft wing tanks. Proposed method consists in that equalising is performed by pumping fuel from aircraft central tank into wing tank with smaller fuel capacity. Note here that valves for feeding fuel to engines and cross valve stay in "shut-off" position to feed fuel from central tank via fuel transfer jet pump into wing tank sump compartment to feed therein an amount of fuel required for balancing fuel amount in both wing tanks.

EFFECT: ruling out unbalance between left and right wing fuel tanks.

2 cl, 1 dwg

FIELD: transportation.

SUBSTANCE: invention refers to aviation, particularly to method of fuel reserve balancing in airplane wing tanks. Method involves balancing by fuel transfer first from central tank and then from the wing tank where fuel amount is larger to the wing tank where fuel amount is less. For that purpose, pump of fuel feed from right wing tank with less fuel to right motor is activated, with fuel feed valves of motors and cross-feed valve in closed position. Fuel is forwarded to active fuel mainline of fuel transfer jet pump from central tank to consumption tank of right wing tank, then fuel is transferred to the right wing tank in amount required to restore balance. Afterwards, if fuel volume in central tank is insufficient to restore balance and fuel is available in wing tanks, fuel is transferred from the left wing tank in ample amount to the right wing tank with smaller fuel reserve. For that purpose, left motor fueling pump is activated in the left wing tank with larger fuel reserve, centralised fuel drain valve is opened, thus connecting left motor fuel feed line with fueling and centralised drain line, with wing tank fueling valves closed. Then right wing tank fueling valve is opened, and fuel is transferred from left to right wing tank until imbalance signal disappears. Fuel transfer from right wing tank with larger fuel reserve to left wing tank with smaller fuel reserve is implemented in the same way.

EFFECT: prevention of transverse imbalance between left and right wing tanks.

4 cl, 1 dwg

FIELD: transport.

SUBSTANCE: invention relates to aircraft engineering, particularly, to method of equalising fuel capacity in aircraft wing tanks. Proposed method consists in that equalising is performed by pumping fuel from aircraft central tank into wing tank with smaller fuel capacity. Note here that with no fuel in central fuel tank, fuel is pumped over from LH wing tank into RH wing tank using fuel priming and centralised drain line of aircraft fuel system, cross feed valve, RH wing tank priming valve and pump to feed fuel to LH engine. Pump to feed fuel to LH engine in LH wing tank is actuated along with centralised fuel drain valve by communicating line of feeding fuel to LH engine with fuel priming and centralised drain line. Then, with closed LH valve, valve to prime RH wing tank is moved to OPEN position to pump fuel over from LH to HT wing tank unless misbalance signal disappears. With no fuel in central fuel tank, fuel is pumped over from RH wing tank into LH wing tank in a similar way.

EFFECT: ruling out unbalance between left and right wing fuel tanks.

4 cl, 1 dwg

FIELD: physics.

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EFFECT: high efficiency of testing characteristics of an aircraft ACT system.

15 cl, 3 dwg

FIELD: transport.

SUBSTANCE: invention relates to aircraft engineering, particularly, to system intended for control over fuel transfer between aircraft fuel tanks. Proposed system includes fuel transfer between aircraft fuel tanks in compliance with preset order dependent upon decrease in aircraft full airborne weight caused by in-flight fuel consumption. Said preset order ensures keeping the aircraft center of gravity in preset range.

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12 cl, 2 dwg

FIELD: transport.

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EFFECT: higher reliability of fuel transfer system and flight safety.

14 cl, 3 dwg

FIELD: transport.

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2 cl, 5 dwg

FIELD: aircraft engineering.

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EFFECT: higher accuracy, validity and efficiency of measurement fuel mass store, fuel reserve supply and control over aircraft balance at standard and emergent conditions of system operation.

1 dwg

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