WO2017018903A1 - Method for placing a payload into orbit using a carrier rocket - Google Patents

Abstract

The invention relates to the field of aerospace engineering and can be used in the production and modernization of various types of rocket systems, including means for placing payloads into near Earth orbit and for placing payloads into a suborbital ballistic trajectory. The essence of the solution consists in a method for placing a payload into orbit using a carrier rocket having a multi-module assembly of combined system rocket modules, said method comprising the following stages: a. during launch of the rocket carrier, liquid propellant sustainer boosters of lateral and central rocket modules are boosted to nominal thrust, b. after longitudinal acceleration is achieved by the carrier rocket, providing for a stable position of the carrier rocket in the trajectory, at least one engine of the liquid propellant sustainer booster of the central rocket module is switched off or the latter is throttled to a level 0.3 lower than the nominal thrust, c. before the liquid propellant sustainer boosters of the lateral rocket modules are switched off, the engine of the liquid propellant sustainer booster of the central rocket module, the thrust of which was reduced earlier, is switched on or throttled to a level 0.3 higher than the nominal thrust, d. the lateral rocket modules are separated off and ejected when the liquid propellant sustainer booster of the central rocket module is switched on, e. the head module is placed into the specified orbit. The technical result is an increase in the payload capacity of carrier rockets in service or under production, with minimal changes to the design thereof.

Description

 METHOD OF ORBITING USEFUL ROCKET LOAD-

CARRIER

Technical field

The invention relates to the field of rocket and space technology and can find application in the creation and modernization of missile systems for various purposes, including means of launching payloads into low Earth orbit and launching payloads to a suborbital ballistic trajectory.

State of the art

The prior art method for launching a payload into orbit with a multifunctional launch vehicle of a combined scheme with marching rocket engines, implemented in the Soviet Union using the Vostok launch vehicle, which is used to launch manned and unmanned spacecraft into low Earth orbit (see. carriers A. V. Alexandrov, V. V. Vladimir, R. D. Dmitriev, S. O. Osipov; Under the general editorship of Prof. S. O. Osipov - M: Military Publishing House, 1981, pp. 19-22 , fig. 1.2). The known method includes attaching, in accordance with the launch program, to the central missile unit of a tandemly located booster missile unit and a head unit with a payload, forming a lower multi-block package of missile units by attaching side missile units to the central missile unit, turning on the start of all marching rocket engine side and central missile blocks, the joint operation of the main and lateral rocket blocks marching liquid propellant rocket engines until the side rocket blocks generate fuel, turning off the neck rocket engine of the side rocket blocks and separation of the side rocket blocks from the central rocket block while continuing to operate the mid-range rocket engine of the central rocket block before fuel is generated from it, shutdown of the main rocket rocket engine of the central rocket block, separation of the tandem located booster rocket and head block from the central rocket block, and the subsequent overclocking the head unit until it goes into orbit. The disadvantage of this scheme is that the central missile unit has large dimensions and mass compared to the side missile units and carries more fuel in its tanks, which ensures a longer operation of its marching rocket engine. This affects the energy capabilities of the packet scheme, since the mass of the central unit includes a large fraction of the mass of the structure that provides fuel storage for the operation of the rocket engine at the stage of flight of the first stage. The closest analogue is the solution "method of putting into payload the orbit of a multifunctional launch vehicle of a combined circuit with marching liquid propellant rocket propulsion systems (jets), a multifunctional launch rocket of a combined circuit with marching rocket engines and a method for its development" described in RF patent jN ° 2161 108, published December 27, 2000. The patent discloses Angara launch vehicles with a packet arrangement of rocket blocks of the first and second stages. In this case, a combined circuit is used with a lower multiblock package of identical rocket blocks having adjustable marching rocket engines with the same nominal thrust, when the launch vehicle starts, the marching rocket engines of the side rocket blocks are brought to rated thrust, and the main marching rocket engine of the central missile block - to thrust equal to 90 ... 100% of the nominal value, and maintain it unchanged until the launch vehicle reaches longitudinal acceleration of 12.7 ... 16.7 m / s (1.3 ... 1.7 g), then reduce the thrust of the mid-range rocket engine missile block up to 0.3 ... 0.5 of the nominal thrust, and after turning off the marching rocket engine of the side rocket blocks, the thrust of the marching rocket engine of the central unit is increased to the nominal value. The use of adjustable marching liquid propellant rocket engines in these missile blocks allows to fully realize the energy capabilities of the lower multiblock package of missile blocks at the start, and a subsequent decrease in the thrust of the main missile propellant rocket engine to 0.3 ... 0.5 of the nominal thrust ensures that it remains in the central rocket fuel supply unit for its marching rocket engine after separation of the side rocket blocks. The thrust reduction of the marching liquid propellant rocket engine of the central missile unit begins after the launch vehicle reaches a longitudinal acceleration of 12.7 ... 16.7 m / s2 (1.3 ... 1.7 g), which ensures a stable position of the launch vehicle on the trajectory . Raising to the nominal value of the thrust of the mid-range rocket engine of the central rocket block after disabling the mid-range rocket engine of the side rocket blocks allows you to fully use the energy capabilities of the central rocket block. The disadvantage of this scheme is that the central missile unit still has a lot of design that provides fuel storage for the operation of the rocket engine at the stage of flight of the first stage. In addition, when throttling to 0.3 ... 0.5 from the nominal thrust, the efficiency of the main engine LRE decreases as a result of a significant decrease in specific impulse. At the same time, the LRE of the central missile unit at this throttle level operate at the maximum level, including experiencing problems in cooling the nozzle, which leads to a decrease in the reliability of the LRE and the launch vehicle as a whole. Technical challenge

The problem to which this invention is directed, is to create a method of flight of launch vehicles with a packet arrangement of rocket blocks of the first and second stages, which allows to increase the mass of the payload. The technical result is to increase the load capacity of the exploited and developed launch vehicles with minimal changes in their design.

Disclosure of invention

To solve this problem, a method for putting payload into orbit with a launch vehicle with a multiblock package of missile units of a combined scheme is proposed, which includes the following steps: a. at the launch of the launch vehicle, marching liquid propellant propulsion systems (LRE) of the lateral and central missile units are launched for nominal thrust,

 b. after the launch vehicle reaches longitudinal acceleration ensuring a stable position of the launch vehicle on the trajectory, at least one engine of the mid-range main propellant rocket engine is switched off or throttled to a level below 0.3 of the nominal thrust,

 c. until the marching rocket engine of the side rocket blocks is turned off, the engine of the marching rocket engine of the central missile block, the thrust of which was previously lowered, is repeatedly turned on or throttled to a level above 0.3 from the rated thrust,

 d. separate and drop side rocket blocks, with the central rocket block LRE turned on,

 e. the head unit is brought out, including the tandemly arranged upper steps to a predetermined path.

When implementing the method, take into account that stage b. occurs when the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s.

To implement the method, they can form a lower multiblock package of missile blocks with non-identical fuel tanks, mass-dimensional characteristics and mid-range rocket engines with the same or different nominal thrust. They can also form a lower multiblock package of rocket blocks with various components of rocket fuel.

To implement the method, they can form a lower multiblock package of missile blocks with a different number of LRE identical LRE with the same nominal thrust.

To implement the method, at the launch of the launch vehicle, the thrust of the marching liquid propellant rocket engine of the central missile unit can be reached and maintained unchanged to a value that ensures that the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s ² . To implement the method, the inclusion of liquid propellant rocket engines can be performed before the marching liquid propellant rocket engines of the side rocket blocks are turned off, sufficient for the central rocket engine to reach the nominal operating mode.

To implement the method, they can install at least one marching propulsion engine rocket engine of the central rocket block, which is turned off or throttled to a level less than 0.3 of the nominal thrust, to move the nozzle nozzles, which are shifted after turning off or throttling the engine to a level less than 0.3 from rated traction.

To implement the method, they can install at least two engines of a mid-range main propellant rocket engine, which shut off or throttle to a level less than 0.3 of the nominal thrust, install a single movable nozzle nozzle, which is shifted after shutting down or throttling the engines to a level less than 0, 3 from rated traction.

Brief Description of the Drawings

FIG. 1 - three-stage launch vehicle with a multiblock package of five identical missile units.

FIG. 2 - two-stage launch vehicle with a multiblock package of five identical missile units.

FIG. 3 - LV with a movable nozzle nozzle on one LRE of the LRE of the central unit.

FIG. 4-PH with a movable nozzle nozzle throughout the LRE of the central unit (a single nozzle nozzle).

FIG. 5 - three-stage launch vehicle with a multi-block package of three missile units with an excellent rocket engine on the central missile unit.

FIG. 6 - three-stage launch vehicle with a multiblock package of three identical missile units.

FIG. 7 - two-stage launch vehicle with a multiblock package of three identical missile units.

FIG. 8 is a flight diagram of a three-stage launch vehicle with a central propellant rocket engine shut down and its subsequent inclusion before separation of the side missile blocks

FIG. 9 is a flight diagram of a three-stage launch vehicle with a central rocket engine rocket engine shut down and then only part of its rocket engine turned on before separation of the side rocket blocks.

In all the figures of the drawings: pos. 1 - central missile unit pos. 2 - lateral missile unit pos. 3 - transition compartment pos. 4 - missile unit of the third stage pos. 5 - head unit pos.6 - LRE rocket of the side rocket block pos.7 - LRS of the central missile block pos.8 - nozzle nozzles for one LRE pos.9 - nozzle single nozzle for the whole LRE of pos.10 - shutter of the head fairing pos.11 - payload

In the figures of drawings 8 and 9, the following designations of the stages of flight are introduced:

A - the launch of the launch vehicle, the inclusion of all LRE LRE of the side and central missile units;

B - shutting down the liquid propellant rocket engine of the central rocket unit, flying with the liquid propellant rocket engine turned off;

In - re-inclusion of liquid propellant rocket engine central missile unit;

G - separation of lateral rocket blocks;

D - separation of the head fairing flaps;

E - separation of the central missile unit with the transition compartment from the tandem located third stage missile unit, the inclusion of the mid-flight rocket engine of the third stage missile unit;

F - exit to orbit, shutdown of the march rocket engine of the third stage rocket block, separation of the payload;

And - extension of the nozzle nozzle of the rocket engine of the central missile unit;

K - re-inclusion of LRE with nozzle LRE of the central missile unit.

The implementation of the invention

The following steps are used to launch a payload into orbit with a launch vehicle with a multiblock package of missile blocks of a combined scheme. Firstly, at the launch of the launch vehicle, marching liquid propellant propulsion systems (LRE) of the lateral and central missile units are brought to nominal thrust.

Secondly, after the launch vehicle reaches longitudinal acceleration, which ensures a stable position of the launch vehicle on the trajectory (this corresponds to an acceleration of 11.8 ... 16.7 m / s ² or 1.2 ... 1.7 g) shutdown of the main propellant rocket engine rocket engine. If the liquid propellant rocket engine consists of several single liquid propellant rocket engines (the first stage of the Falcon rocket), then not all liquid propellant rockets can be turned off, but only a part of them. This is similar to the throttling of the LRE of the central block of the Angara launch vehicle. Technically, each individual engine has the ability to autonomously control traction and turn on / off, which is implemented in almost all launch vehicles. When the entire rocket engine shuts down, then commands are issued to all engines at once. It should be noted that engines such as 11D58 (booster block type "DM"), S.98M (booster block type "breeze") and other engines of booster blocks have the ability to re-enable in flight. Also, the Merlin-ID engine on the Falcon-9 LV has the ability to re-enable the flight, which it implements when launching the spent missile unit. For such engines, if they are restarted, either several ampoules with starting fuel are used, or tanks with starting fuel from which it is supplied in the doses necessary to start the engine. Electric ignition may be used. One of its varieties is laser ignition (patent application of the Russian Federation No. 2012157504 company Spectralazer, published on 10.07.2014). In this case, it may be necessary to refine the design itself for re-inclusion. Following the example of the Falcon-9 launch vehicle, these improvements are minor and relatively easy to implement. As will be shown below in the implementation examples, in order to increase the mass of cargo put into orbit with a fixed mass of a rocket with fuel, the rocket engine of the central missile unit should be turned off during the launch of the rocket into orbit, but a positive effect, though to a lesser extent, will be in the case of a decrease in the thrust of the main rocket engine central missile unit to a non-zero level, but below 0.3, i.e. in case of throttling to a level below 0.3. The liquid propellant rocket shutdown can occur somewhat earlier or later than the moment the launch vehicle reaches acceleration 11, 8 ... 16.7. These values can vary significantly and depend on the specific rocket. Thirdly, before the marching rocket engines of the side missile blocks are turned off, the marching rocket engines of the central missile block are switched on again or throttled to a level above 0.3.

Finally, at the end, the lateral rocket blocks are separated and dropped, with the central rocket block rocket engine turned on, and after the end of the working fuel reserves in the central rocket block, it is separated from the rocket block with the head block with the subsequent dispersal of the head block before it enters a predetermined orbit.

To implement the method, they can form a lower multiblock package of missile blocks with non-identical fuel tanks, mass-dimensional characteristics and mid-range rocket engines with the same or different nominal thrust. They can also form a lower multiblock package of rocket blocks with various components of rocket fuel. To implement the method, they can form a lower multiblock package of missile blocks with a different number of LRE identical LRE with the same nominal thrust. To implement the method, at the launch of the launch vehicle, the thrust of the marching liquid propellant rocket engine of the central missile unit can be reached and maintained unchanged to a value that ensures that the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s. The specific thrust value is found by calculation or experimentally for a specific rocket model, which provides the maximum payload mass and design restrictions.

To implement the method, the inclusion of liquid propellant rocket engines can be performed before the marching liquid propellant rocket engines of the side rocket units are disconnected for a time sufficient for the central rocket engine to reach the nominal operating mode. This time is determined by the dynamic characteristics of each particular engine. The range can theoretically be any, although from the conditions of efficiency it can be from 0 to 100 seconds. The inclusion of liquid propellant rocket engines can be performed after separation of the side rocket blocks, i.e. with the engines off. This may be due to the fact that when separating missile blocks, requirements for reducing loads, including aerodynamic ones, may arise. That is, in order to reduce the load during separation, it is possible to share with the turned off engines of the central unit.

To implement the method, at least one engine marching liquid propellant rocket engine of the central missile unit, which is turned off or throttled to a level less than 0.3, can install a movable nozzle nozzle, which is shifted after turning off or throttling the engine to a level of less than 0.3. Nozzle nozzles can further increase the mass of the payload. On the Angara launch vehicle, such an opportunity was considered and an effect was obtained. But the nozzle nozzles on the engine, shifted at the start, do not allow the rocket to exit the launch pad, and it is not possible to shift the nozzles during the flight on a running engine due to gas dynamics problems. In the proposed solution, in the case of the engine turned off or throttling close to zero, it turns out to shift the nozzles during the flight.

To reduce the cost of the process, at least two engines of a mid-range main propellant rocket engine, which shut off or throttle to a level of less than 0.3, can install a single movable nozzle nozzle, which is shifted after turning off or throttling the engines to a level of less than 0, 3.

 Implementation example

 We will consider the result of the proposed technical solution by the example of modernization of the 3-stage Angara-5 launch vehicle, in which the 1st and 2nd stages are a package assembled from 5 standardized missile blocks (1 central unit and 4 side blocks ) The basic input data adopted in the calculations are presented in table 1.

 Table 1 - Basic input data.

 Name Value

 Specific impulse of thrust of rocket engine engines in

 void in nominal mode (coefficient

throttling to _d = 1,0) R _{beats 1; 0} , kgf · s / kg: 337.5

 Specific impulse of thrust of rocket engine engines in

void in maximum throttling mode (k _d =

0.3) R _beats about, s, kgf * s / kg: 330.8

 Specific impulse of thrust of rocket engine engines in

void in nominal mode (k _d = 1,0) with nozzle

nozzle R _{beats 1; 0 n} , kgf · s / kg: 343.5 ... 358.7

Gross mass of one missile block M _0r b, t 138.5

 The working fuel supply of one rocket block Μη, β, t 127.5

Initial mass of the 3rd stage M ₀₃ , t 80.5

The ratio of the operating time of the central engines missile block before throttling (k _d = 1,0) k

the entire operating time of engines of the 1st stage, t _d : 0.2

The existing flight scheme of the Angara-5 launch vehicle at the 1st and 2nd stage sections provides for starting all the engines of the central and side blocks at the start at nominal speed, when the throttle coefficient is to _d = 1.0. Upon reaching the launch vehicle longitudinal acceleration of 1,6 g, the central unit engines are throttled to a level to _d = 0.3, the lateral blocks motors remain in nominal mode. After completion of the flight stage of the 1st stage and separation of the side blocks, the engines of the central block are again brought back to the nominal mode.

 The mass of fuel of the central unit consumed in the flight section of the 1st stage, when the engines are operating in nominal mode:

M _t12 ι, ο = M _tr b "Td = 127.5 · 0.2 = 25.5 t.

The mass of fuel of the central unit spent on the flight section of the 1st stage, when the engines are operating in throttle mode with k _d = 0.3:

_T12 M _o, s = (M _TP6 - _{T12 1> 0)} · a _a - t = 31.2.

 The mass of fuel of the central unit spent on the entire flight section of the 1st stage:

t ₁₂ = M _{t12 1 0} + M _t12 o s = 25.5 + 31.2 = 56.7 t

 The total fuel mass of 4 lateral and 1 central blocks, consumed throughout the flight section of the 1st stage:

t ₁ = 4 · _tr6 + M _t12 = 4 · 127.5 + 56.7 = 566.7 t.

 The fuel mass of the central unit remaining for the flight section of the 2nd stage:

M = M _{T2 TRB} - M _r12 = 127.5 - 56.7 = 70.8 m.

 Launch mass of the entire launch vehicle:

0 ₁ = 5 - _0р6 + ₀₃ = 5 · 138.5 + 80.5 = 773.0 t.

 The initial mass of the entire launch vehicle in the area of the 2nd stage:

M ₀₂ = M _0rb - _t12 + M ₀₃ = 138.5 - 56.7 + 80.5 = 162.3 t.

The average specific thrust impulse of the engines of the central and side rocket blocks in the void on the flight section of the 1st stage, taking into account the throttling of the central block to the level k _d = 0.3:

S ^ P dt P _rb - t _x · [4 + t _d + (1 - t _d ) - / Cd]

/ "m dt (R _b / R _bpm , o) · · [4 + t _d + (1 - t _d ) · / s _d · R _{bpm 1> 0} / P _{bpm 0} ] ₃ ] Ore ι, ο · [4 + t _d + (1 - t _d ) · k _d ] 337.5 - [4 + 0.2 + (1 - 0.2) - 0.3]

 - = 337.1 s

4 + t _d + (1 - t _d ) - k _d · R _y d 1, o / Rud oz + 0.2 + (1 - 0.2) · 0.3 - 337.5 / 330.8

 where t] is the total operating time of all engines of the 1st stage.

 According to the Tsiolkovsky formula, the characteristic speed that the launch vehicle picks up on the flight section of the 1st and 2nd stages:

M _t1 \ _. . Μ ,.

them. ap = -Py, D1 ^{50 '1n} ( ¹ - ^) - ^p ^ - ⁵ ° - ¹ⁿ ( ¹ - ^) ⁼

 / 566.7 \ / 70.8 \

 = -337.1 · 9.81 - In - - 337.5 · 9.81 · In - = 6265.8 m / s where go is the acceleration of gravity.

 Next, we evaluate the result that we will get when using the proposed technical solution, if instead of throttling the engines of the central unit we completely turn them off. Then the mass of fuel of the central unit spent on the entire flight section of the 1st stage:

t ₁₂ = Af _Tl2 i, o = 25.5 t.

 The total fuel mass of 4 lateral and 1 central blocks, consumed throughout the flight section of the 1st stage:

t ₁ = 4 · M _trb + M _t12 = 4 · 127.5 -I- 25.5 = 535.5 t.

 The fuel mass of the central unit remaining for the flight section of the 2nd stage:

M = M _{T2 TRB} - _T12 = 127.5 - 25.5 = 102.0 m.

We calculate the initial mass of the 3rd stage, which will correspond to the same characteristic speed that the booster gains in the flight section of the 1st and 2nd stages when throttling the engines of the central unit (to _d = 0.3). We write the Tsiolkovsky formula for the flight section of the 2nd and 2nd steps in the form:

 Consider that:

0 ₁ = 5 · _0rb + M ₀₃ and M ₀₂ = _0rb - M _t12 + M ₀₃ , then the Tsiolkovsky formula will take the form: 5 - _0rb + ₀₃ y _0rb - _t12 + M ₀₃

Thus, we have obtained a nonlinear equation for the initial mass of the 3rd stage M ₀₃ . In the general case, it can be solved by numerical methods. We will take into account that when the central unit engines are turned off, we have no loss of specific thrust impulse as during throttling, therefore, the specific impulse of thrust in the void in the flight section of the 1st stage is the same as in the segment of the flight of the 2nd stage, and is equal to the specific impulse of thrust in the void in the nominal mode P _beats . Then the equation for the initial mass is 3- the first stage of Moz can be rewritten in the form:

 and reduce to the usual quadratic equation of the form:

a - M ₀₃ ² + b - M ₀₃ + c = 0

 Where, a = E - 1

B = (6 · _0rb - _t12 ) - E - 6 · M _0rb + M _t1 + M _tr6

s = 5 · _0rb · (M _0r6 - M _t12 ) - E - (5 · _0rb - M _t1 ) · (M _0rb - M _trb )

 and after substitution of values:

 6265.8 \

 E = exp [-————) = 0.150696

 \ 9.81 - 337.5 /

 a = 0.150696 - 1 = -0.849304

 B = (6 · 138.5 - 25.5) - 0.150696 - 6 - 138.5 + 535.5 + 127.5 = -46.614365 s = 5 · 138.5 · (138.5 - 25 , 5) - 0.150696 - (5 - 138.5 - 535.5) - (138.5 - 127.5) =

 = 10065,339456

 Solving this quadratic equation, we consider only its positive root:

——— Vfe ² - 4ac

0 _{3 off} = 84.83 tons

 2a

Thus, turning off the engines of the central unit compared to their throttling to the level k _d = 0.3 allows you to increase the initial mass of the 3rd stage by: · 100% = 5.4%

The payload mass of the 3rd stage M _mon , which is the payload of the entire launch vehicle, is associated with the initial mass of the 3rd stage M _{03 by the} ratio:

M _mon = μ _ΠΗ3 - M _03.

where μ „ _Η 3 is the relative mass of the 3rd stage payload, which depends on the characteristic speed developed in the flight section of the 3rd stage and the mass perfection of the 3rd stage design. The trajectory parameters do not change; therefore, we consider the value of the characteristic velocity unchanged. Mass excellence structures while increasing the absolute mass of the structure only improves, so the ability to increase the starting mass of the 3rd stage by 5.4% will increase the payload mass by at least 5.4%, which will be 1 on the scale of the Angara-5 launch vehicle , 3 t.

Now we will evaluate the effect of using movable nozzle nozzles on the engine chambers of the central unit. At the time of the launch of the launch vehicle, the nozzle nozzles were removed to provide maximum earth thrust for the engines of the central unit. In the flight section of the 1st stage, after turning off the engines of the central unit, it becomes possible to safely extend the nozzle nozzles, so that when the engines of the central unit are switched back on, which occurs after separation of the side blocks, i.e. at high altitudes, provide maximum void traction. According to various estimates, the specific thrust impulse of engines with a nozzle nozzle in a void can range from 343.5 to 358.7 kgf · s / kg (see table 1). We substitute these values into equation (1) as P _yD 2. Then, solving equation (1) by numerical methods, we obtain the following expected range of increase in the initial mass of the 3rd stage when the central unit engines are turned off together with the use of nozzle nozzles in the flight section of the 2nd stage:

 that on the scale of the Angara-5 launch vehicle will be 1.8 ... 3.0 tons.

 Thus, the use of movable nozzle nozzles in the flight section of the 2nd stage can, in addition to turning off the engines of the central block in the flight section of the 1st stage (Δο ™, ^ 5.4%), increase in the payload of 1.9 ... 6 ,8 %.

 We examined the result of the proposed technical solution as an example of the modernization of a specific product. We now show this with the help of theoretical calculations.

The relative mass of the payload of the first 2 stages of the launch vehicle: mon ^{--— ~ mon2} _— A * mon1 "^ mon2 g K *> -) where μ _ΠΗΐ = is the relative mass of the payload of the 1st stage, μ _ΠΗ2 = —— = —— - the relative mass of the payload of the 2nd stage,

Stage 1 payload mass: Μ _ΠΗ Ι = Afoi - Μ _τ1 - m _TOll - t _dv11 - m _npll (3) where Μοι is the initial (starting) mass of the 1st stage,

M _T1 - the total mass of fuel of the side and central blocks spent on the entire flight section of the 1st stage,

 πΐτοπ - mass of fuel compartments of the side blocks of the 1st stage,

 Shdvp - the mass of propulsion systems of the side blocks of the 1st stage,

 Shprp - the mass of the control system and all other compartments of the side blocks of the 1st stage.

The total mass of fuel of the side and central blocks consumed throughout the flight section of the 1st stage M _t1 can be expressed through its starting mass:

Mn = ₀₁ - (1 - Μκΐ) (4) where μ _κ ι is the relative final mass of the first step.

The mass of fuel of the central unit spent on the entire flight section of the 1st stage:

where is the total operating time of all engines of the 1st stage,

t ₁₂ - second fuel consumption of the engines of the Central unit,

Pi ₂ - thrust of the engines of the Central unit,

Ru _{d 12} - specific impulse of thrust of the engines of the central unit.

 We introduce the designation of the relative starting thrust of the engines of the central unit:

Ρ - - (6) where Po is the starting thrust of the engines of the central unit,

P ₀₁ - starting thrust of all engines of the 1st stage.

Then the thrust of the engines of the central unit can be written as:

Ρ ₁₂ = Ροι · Ρ · (7) where k _d is the current throttle coefficient of the engines of the central unit.

 Given the expression (7), the formula for the fuel mass of the central unit (5) will take the form:

where is the total operating time of all engines of the 1st stage, R _beats ι is the average integral specific impulse of thrust of all engines of the 1st stage. Then the expression for the mass of fuel of the central unit consumed throughout the flight section of the 1st stage will take the form:

M _t12 = M _t1 - P - / s _{d cf} (8) where k _{d cf} is the average integral throttle coefficient of the engines of the central unit:

 We will describe the masses of the structural elements of rocket blocks using simplified mass coefficients. Then, taking into account expressions (4) and (8):

™ _Τ οΐι = -l - Μ _τ12 ) · а _r = M _Tl (l - Р · к _{л Wed} ) · _r = ₀₁ (1 - μ _κ ι) (ΐ - Р '^ _Д с _Р )''ι ( J) where αι is the mass coefficient of the fuel compartments of the side blocks of the 1st stage (the ratio of the mass of the fuel compartments to the mass of fuel in these compartments).

 The mass of the engines of the side blocks of the 1st stage, taking into account the expression (7):

™ _Д нн = 7ι · Roy = Υι · (οι - Р012) = Υι "^ oi · (1 - P) (I) where i is the mass coefficient of the engines of the side blocks of the 1st stage (the ratio of the mass of the engines to the starting thrust of these engines Ροπ) ·

Starting thrust of all engines of the 1st stage can be represented as:

Ρ „ι = n ₀₁ · ₀₁ (12) where n ₀₁ is the starting overload of the 1st stage.

We substitute (12) into (11), then:

W = 7ι · Po1 - ₀₁ · (1 - P) (13) The mass of the control system and all other compartments of the side blocks of the 1st stage:

Sharp = βι · ₀₁ (14) where βι is the mass coefficient of the control system and all other compartments (the ratio of the mass of the control system and all other compartments of the side blocks of the 1st stage to the starting mass of the 1st stage).

 We substitute the expressions (4), (10), (13) and (14) in (3):

Μ _ΠΗ Ι = M ₀ 1 - M ₀₁ (1 - _κ ι) - M ₀₁ (1 - μ _κί ) (ΐ - P · c _{d av} ) · a _d - ιΠ ₀ ι ^ οι (ΐ ~ P) ~ βι ^Μ οι

с_Р«1 - Χιη₀ι(ΐ - Ρ) - β₁ == Ρ · к_{д ср} (1 - μ_κ1) + κ₁(1 + α ) - _7ι ^η ₀₁(ΐ - Ρ) - βι (15)= ₀₁ - [μ _κ1 - (1 - μ _κ1 ) (ΐ - P · / Cd cf.) - "ι - i oi (l ~ P) ~ βχ] Then the relative mass of the 1st stage payload: βπΗΐ = μ _κ ι - (1 - Μκΐ) (ΐ - Ρ · cp) · ι - 7in ₀ i (l - Ρ) ^~ βι = μ _κ ι ^~ "ι + Ρ '

with _P «1 - Χιη ₀ ι (ΐ - Ρ) - β ₁ == к · _{d cp} (1 - μ _κ1 ) + κ ₁ (1 + α) - _7ι ^η ₀₁ (ΐ - Ρ) - βι (15 )

Denote:

Α ^ Ρ - α ^ Ι - μ ^) (16) Β = _κ ι (1 + «ι) - 7ιη ₀₁ (ΐ - Ρ) - βι (17) then

 Similarly to (3) we write the mass of the payload of the 2nd stage:

M '= M _{H2 02} - _m2 - mn _TO2 - _dB2 t - t _np2 (19) wherein M ₀₂ - initial weight of 2nd stage,

M is the mass of fuel _T2 of the central unit, expended in the area of flight stage 2,

W _T0 2 - the mass of the fuel compartment of the Central unit (2nd stage),

 tdv2 - mass of the propulsion system of the central unit (2nd stage),

w _P p2 - the mass of the control system and all other compartments of the central unit (2nd stage).

The mass of fuel of the central unit consumed in the flight section of the 2nd stage M _t2 can be expressed in terms of the initial mass of the 2nd stage:

Μ _τ2 = Μ ₀₂ · (1 - μ _Κ 2) (20) where μκ2 is the relative final mass of the second stage.

 The mass of the fuel compartment of the central unit is written taking into account the fact that part of the fuel of the central unit is also consumed in the flight section of the 1st stage. Then, taking into account expressions (20), (8) and (4):

™ _T0 2 = ^(m m2 + Μ _τ12) · ₂ = M ₀₂ (1 - μ _κ2) · ₂ + M ₀₁ (1 - μ _κ1) · P · to _{d cp} and ₂ (21) where a ₂ - mass ratio fuel compartment of the central unit (the ratio of the mass of the fuel compartment to the mass of fuel in this compartment).

 The mass of the engines of the central unit (2nd stage), taking into account expressions (7) and (12):

™ _dv2 = _Vg · Roi = Vg 'Poi' P = V g ¹ "oi ¹ M ₀₁ - P (22) where y2 is the mass coefficient of the engines of the central unit (the ratio of the mass of the engines of the 2nd stage to the starting thrust of these engines P ₀₁₂ ) .

 The mass of the control system and all other compartments of the side blocks of the 2nd stage:

где β₂ - массовый коэффициент системы управления и всех прочих отсеков (отношение массы системы управления и всех прочих отсеков центрального блока к начальной массе 2-й ступени).

where β ₂ is the mass coefficient of the control system and all other compartments (the ratio of the mass of the control system and all other compartments of the central unit to the initial mass of the 2nd stage).

 We substitute the expressions (20) - (23) in (19):

т.е.: M _PN 2 = M ₀₂ - M ₀₂ (1 - μ _κ2 ) - M ₀₂ (1 - _κ2 ) 2 - Af ₀₁ (l -

those.:

(24)

 Consider that the initial mass of the 2nd stage is the payload of the 1st stage:

M ₀₂ = M _PN1 = μ _ΠΗΐ - M ₀ 1 (25) then

/ = ΓΓ ^ = Mkg (1 + "2) -— [(1 - with _P " 2 ~ Y2 1] ~ "r ^~ βι (26) m02 / * mon

 Thus, from (2), taking into account (26), the relative mass of the payload of the first two stages of the launch vehicle:

Mpn = Μ _ΠΗ Ι - [μ _Κ 2 (1 + "r) -α ₂ -β ₂ ] -Ρ · (1- μ _κ1 ) Λ _{Λ cf 2} - Р · γ ₂ ^η ₀₁ (27) We denote:

L ₂ = μ _κ2 (1 + а ₂ ) - ₂ - β ₂ (28)

Β ₂ = Ρ · {1-μ _κ1 ) α ₂ (29)

C ₂ = Py ₂ n ₀ i (30) Then from (27), taking into account (18):

mon = (cp ^ i + Bi) - ^ 2 - # 2 - k _AC p - C ₂ = А А ₂ · s _d cp + # 1 ^ 2 - B ₂ · / s _d c _P - C ₂ (31) To assess the influence of the average integral throttle coefficient of the central unit engines k _{d av} on the relative payload mass of the first 2 stages of the launch vehicle μ _ΠΗ , we consider from (31) its partial derivative taking into account (16), (28) and (29):

 9μ »

P (1 - μ _κ1 ) · « _! - [μ _κ2 (1 + α ₂ ) - α ₂ -? ₂ ] - P (l - μ _κ ι) ¹ «2 (32)

S / Wed Wed

Find the extremum of the function μ _ΠΗ ⁼ f (cdsd)

—— = Ρ (1-μ _κ1 ) -α ₁ μ _κ2 (1 + α ₂ ) - α ₂ -? ₂ - = 0

we are interested in the case when:

μ _κ2 (1 + α ₂ ) -α ₂ - /? ₂ - ^ = 0 (33)

"one We determine the value of the relative final mass of the 2nd stage corresponding to the extremum point:

α ₂ / α ₁ + α ₂ + β ₂

μ _Κ 2 =: (34)

^ ^κΖ 1 + a ₂

Consider the left and right sides of expression (34) in relation to the existing values of variables for a launch vehicle, in which the central and side blocks use the same components of rocket fuel. Then a ₂ "i.e. α ₂ Λ * ι) »1 and, therefore, given that all mass coefficients are positive, the right-hand side of expression (34):

At the same time, the left side of expression (34) μ _κ2 <1, because the final mass of the 2nd stage is always less than the initial. Thus, for a launch vehicle in which the central and side blocks use the same components of rocket fuel, expression (34) must be written as inequality:

and _g / a _x + α ₂ + β ₂

μ _Κ 2 <: (36)

^ ^κ2 1 + α ₂

This suggests that the partial derivative δμ _ΠΗ / δ1ί _{Ά (} . _Ρ <0, i.e., to increase the relative payload mass μ _π „of the first 2 stages of the launch vehicle, in which the central and side blocks use the same and the same propellants, it is necessary to reduce the mean-rate throttling engines central unit to _{cf. d} up to their complete failure. However, disabling motor of the central unit is only possible when the stable motion of the trajectory, so in practice to _d always _cf. udet greater than zero, otherwise lost the basic idea of the package - creation of an additional launch rocket thruster unit 2nd stage.

Let us estimate the value of the ratio of the mass coefficients of the fuel compartments of the central and side blocks corresponding to the extremum point, after which the partial derivative δμ _ΠΗ / δ! Ί _{Ά (} . _Ρ > 0, and to increase the payload mass μ _ΠΗ , on the contrary, it is necessary to increase the average _integral throttle coefficient of the engines of the central block k _{d cf.} For this, we write expression (33) in the form of the inequality:

 a2

μ _κ2 (1 + α ₂ ) - a ₂ - β ₂ -—> 0

 "one

or: - <μ _κ2 + _Κ 2 · «2 -« 2 - βζ (37) we write with respect to α ₂ :

α ₂ - «ι - _Κ 2 · α ₂ +« ι '«2 <« ι - μ _κ2 - <* ι · βζ

or:

α ₂ - [1 + « _! · (1 - μ _κ2 )] <« _! - (μ _κ2 - /? ₂ )

and finally:

- · [1 + " _! · (1 - μ _κ2 )] <μ _κ2 - β ₂ (38)

Let us evaluate expression (38), taking into account the fact that for existing structures, the variables in this expression can take the following values:

 g = 0.05 ... 0.10

μ _κ2 = 0.3 ... 0.6,

β ₂ = 0.01 ... 0.02

Then the value of [1 + · (1 - μ _κ2 )] does not exceed 1.07. Thus, taking into account the range of values of β ₂ , we can conclude that to stop the growth of the relative mass of the payload of the first 2 stages of the launch vehicle μ _ΠΗ with a decrease in the average _integral throttle coefficient of the engines of the central unit to _{q Wed} , it is necessary that the ratio the mass coefficients of the fuel compartments of the central and side blocks (a ₂ / a,) did not exceed the value of the relative final mass of the 2nd stage, i.e.:

 or, in other words, with the same fuel reserves, the tanks of the central unit should be 2 ... 3 times lighter than the tanks of the side unit, which is practically not feasible.

Claims

Claim 1. A method of launching a payload into orbit with a launch vehicle with a multiblock package of missile blocks of a combined circuit comprising the following steps: a. at the launch of the launch vehicle, marching liquid propellant propulsion systems (LRE) of the lateral and central missile units are launched for nominal thrust, b. after the launch vehicle reaches longitudinal acceleration ensuring a stable position of the launch vehicle on the trajectory, at least one engine of the mid-range main propellant rocket engine is switched off or throttled to a level below 0.3 of the nominal thrust, c. until the marching rocket engine of the side rocket blocks is turned off, the engine of the marching rocket engine of the central rocket block, the thrust of which was previously lowered, is repeatedly turned on or throttled to a level above 0.3 from the rated thrust d. separate and drop side rocket blocks, with the central rocket block LRE turned on, e. the head unit is brought out, including the tandemly located upper steps to a predetermined path. 2. The method according to p. 1, characterized in that stage b. occurs when the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s. 3. The method according to p. 1, characterized in that they form a lower multiblock package of rocket blocks with non-identical fuel tanks, mass-size characteristics and mid-range rocket engines with the same or different nominal thrust. 4. The method according to p. 1, characterized in that they form a lower multiblock package of rocket blocks with various components of rocket fuel. 5. The method according to p. 1, characterized in that they form a lower multiblock package of missile blocks with a different number of LRE identical LRE with the same nominal thrust. 6. The method according to p. 1, characterized in that, at the launch of the launch vehicle, the thrust of the marching liquid propellant rocket engine of the central missile unit is output and maintained constant to a value that ensures that the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s ² . 7. The method according to p. 1, characterized in that the inclusion of liquid propellant rocket engines is performed before the marching rocket propellant rocket engines of the side rocket units are turned off, sufficient for the central rocket engine to reach its nominal operating mode. 8. The method according to p. 1, characterized in that at least one marching propulsion engine of the central rocket engine, which is turned off or throttled to a level less than 0.3 of the nominal thrust, sets a movable nozzle nozzle, which is shifted after shutdown or throttling the engine to less than 0.3 of the rated thrust. 9. The method according to p. 1, characterized in that at least two engines of the marching liquid propellant rocket engine of the central rocket block, which shut off or throttle to a level less than 0.3 of the nominal thrust, install a single movable nozzle nozzle, which is shifted after shutdown or throttling engines to less than 0.3 of the rated thrust.