RU2595092C1 - Method for payload orbital injection by carrier rocket - Google Patents

Abstract

FIELD: rocket equipment; astronautics.

SUBSTANCE: invention relates to aerospace engineering. Method for payload orbital injection by carrier rocket with a multi-block pack of rocket pods of combined scheme includes multiple stages. At start of carrier rocket, cruise liquid jet propulsion devices (LJPD) of lateral and central rocket units are released at nominal thrust. After carrier rocket achieves longitudinal acceleration, which provides stable position of carrier rocket on trajectory, turning off at least one engine of cruse LJPD of central rocket unit is performed or, it is throttled to below 0.3 of nominal thrust. To disconnect cruse LJPD of lateral rocket units, engine or cruise LJPD of central rocket unit is repeatedly turned on or is throttled to a level above 0.3 of nominal thrust, thrust of which was previously decreased. Separating and releasing lateral rocket units when LJPD of central rocket unit is turned on. Head unit is brought to preset orbit.

EFFECT: increased carrying capacity of operated and created carrier rockets with minimal changes in their design.

9 cl, 9 dwg, 1 tbl

Description

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 V.A. Aleksandrov, V.V. Vladimirov, R.D. Dmitriev, S.O. Osipov; Under the general editorship of Prof. S.O. Osipov - Moscow: Military Publishing House, 1981, p. 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 the payload into orbit with a multifunctional launch vehicle of a combined scheme with marching liquid propellant rocket propulsion systems (LRE), a multifunctional launch vehicle of a combined scheme with marching liquid propellant rocket engines and a method for its development" described in RF patent No. 2161108, published December 27, 2000. The patent discloses launch vehicles of the Angara family 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 keep it unchanged until the launch vehicle reaches longitudinal acceleration of 12.7 ... 16.7 m / s ² (1.3 ... 1.7 g), then reduce the thrust of the main rocket engine of the central missile unit to 0, 3 ... 0.5 of rated 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 the indicated missile blocks allows to fully realize the energy capabilities of the lower multiblock package of missile blocks at the start, and the subsequent decrease in the thrust of the main rocket propellant rocket engine to 0.3 ... 0.5 of the nominal thrust ensures that the reserve in the central missile block fuel 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 rocket block begins after the launch vehicle reaches longitudinal acceleration of 12.7 ... 16.7 m / s ² (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 due to 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 reduces 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 is proposed for launching a payload into orbit with a launch vehicle with a multiblock package of missile blocks of a combined scheme, 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 has reached 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 the 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.

When implementing the method, consider that stage b. occurs when the launch vehicle reaches longitudinal acceleration of 11.8 ... 16.7 m / s ² .

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 main engine rocket engine of the central rocket block, which is shut off or throttled to a level less than 0.3 of the nominal thrust, to install a movable nozzle nozzle, which is shifted after turning off or throttling the engine to a level less than 0.3 from rated traction.

To implement the method, at least two engines of a mid-range main propellant rocket engine, which shut off or throttle to a level less than 0.3 from the nominal thrust, can install a single movable nozzle nozzle that is shifted after turning off or throttle 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 - LV with a movable nozzle nozzle throughout the LRE of the central unit (a single nozzle nozzle).

FIG. 5 - a three-stage launch vehicle with a multiblock 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 shutdown with its subsequent inclusion before separation of the side missile blocks

FIG. 9 is a flight diagram of a three-stage rocket launcher with a central propellant rocket engine shutdown with the subsequent inclusion of only part of its rocket engine before separation of the side missile blocks.

In all figures of the drawings:

pos. 1 - central missile unit;

pos. 2 - side rocket block;

pos. 3 - transition compartment;

pos. 4 - missile block of the third stage;

pos. 5 - head unit;

pos. 6 - rocket engine side rocket unit;

pos. 7 - LREU central missile unit;

pos. 8 - nozzle nozzles on one rocket engine;

pos. 9 - nozzle single nozzles for the entire rocket engine;

pos. 10 - sash 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;

Ε - 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, ensuring 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), the rocket engine is switched off central missile unit. 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 enable / disable, 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-1D 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 switched back on, 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.

Shutting down the liquid propellant rocket engine can occur 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 stocks of fuel 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 given 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 ² . The specific thrust value is found by calculation or experimentally for a specific rocket model, which provides the maximum payload mass and design constraints.

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 of the mid-range main propellant rocket engine, which is turned off or throttled to a level of less than 0.3, can install a movable nozzle nozzle, which is shifted after turning off or throttling the engine to 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.

The existing flight scheme of the Angara-5 launch vehicle in 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 k _d = 1.0. After the launch vehicle reaches a longitudinal acceleration of 1.6 g, the engines of the central unit are throttled to the level k _d = 0.3, while the engines of the side blocks remain at their nominal conditions. 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 1.0} = M _trb · τ _d = 127.5 · 0.2 = 25.5 t.

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

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

M _t12 = M _{t12 1.0} + M _{t12 03} = 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:

M _t1 = 4 · M _trb + 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 _r2 = -M M _{TRB r12} = 127,5-56,7 = 70.8 m.

Launch mass of the entire launch vehicle:

M ₀₁ = 5 · M _0rb + M 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 -M _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:

where t ₁ - 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:

where g _o 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:

M _t12 = Mt _{12 1.0} = 25.5 t.

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

M _t1 = 4 · M _trb + M _t12 = 4 · 127.5 + 25.5 = 535.5 t.

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

M _r2 = -M M _TRB 127,5-25,5 _r12 = m = 102.0.

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 (k _d = 0.3). We write the formula of Tsiolkovsky for the flight section of the 1st and 2nd stages in the form:

Consider that:

M ₀₁ = 5 · M _0rb + M ₀₃ and M ₀₂ = M _0rb- M _t12 + M ₀₃ ,

then the Tsiolkovsky formula will take the form:

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 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 thrust impulse in the void in the 1st stage flight section is the same as in the 2nd stage flight section, and is equal to the specific thrust momentum in the void at the nominal mode P _{beats 10} . Then the equation for the initial mass of the 3rd stage M ₀₃ can be rewritten in the form:

and reduce to the usual quadratic equation of the form:

a · M ₀₃ ² + b · M ₀₃ + s = 0

Where, a = E-1

b = (6 · M _0rb- M _t12 ) · E-6 · M _0rb + M _t1 + M _trb

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

$$E=\exp \left(-\frac{{\nu }_{x\mathrm{but}R}}{g_0{P}_{\mathrm{at}\mathrm{<figure-callout\; id="1"\; label="d"\; filenames="00000063.png,00000069.png"\; state="\{\{state\}\}">d</figure-callout>}\mathrm{1,0}}}\right)$$

and after substitution of values:

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:

Thus, turning off the engines of the central unit in comparison with their throttling to the level k _d = 0.3 allows you to increase the initial mass of the 3rd stage by:

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 _pn = μ _pn3 · M ₀₃ ,

where µ _PN3 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. The massive perfection of the structure with an increase in 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 is on the scale of the Angara-5 launch vehicle will be 1.3 tons

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 _beats2 . 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 the nozzle head to be shifted in the flight portion stage 2 can give, in addition to switching off the engine of the central unit in the area of the 1st stage of flight (Δ = 5.4% _none) payload gain 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 carrier rocket:

- относительная масса полезной нагрузки 1-й ступени,Where

- the relative mass of the payload of the 1st stage,

- относительная масса полезной нагрузки 2-й ступени,

- the relative mass of the payload of the 2nd stage,

Stage 1 payload mass:

where M ₀₁ is the initial (starting) mass of the 1st stage,

M _t1 is the total fuel mass of the side and central blocks consumed throughout the flight section of the 1st stage,

m _to11 - the mass of the fuel compartments of the side blocks of the 1st stage,

m _dv11 - the mass of the propulsion systems of the side blocks of the 1st stage,

m _pr11 - 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:

where µ _k1 is the relative final mass of the 1st stage.

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

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

- секундный расход топлива двигателей центрального блока, $${\dot{m}}_{12}$$

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

P ₁₂ - thrust of the engines of the Central unit,

R _ud12 - 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:

where R ₀₁₂ - 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:

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 t ₁ - the total operating time of all engines of the 1st stage,

R _{beats 1} - 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:

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):

$$m_{t\mathrm{about}\mathrm{eleven}}=\left(M_{t\mathrm{one}}-M_{t12}\right)\cdot {\alpha }_{\mathrm{one}}=M_{t\mathrm{one}}\left(\mathrm{one}-\overline{R}\cdot k_{d\mathrm{from}R}\right)\cdot {\alpha }_{\mathrm{one}}=M_{01}\left(\mathrm{one}-{\mu }_{\mathrm{to}\mathrm{one}}\right)\left(\mathrm{one}-\overline{R}\cdot k_{d\mathrm{from}R}\right)\cdot {\alpha }_{\mathrm{one}}\text{(10)}$$

where α ₁ 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):

where γ ₁ 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 is P ₀₁₁ ).

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

where n ₀₁ is the starting overload of the 1st stage.

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

The mass of the control system and all other compartments of the side blocks of the 1st stage:

where β ₁ 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):

Then the relative mass of the payload of the 1st stage:

Denote:

then

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

where M ₀₂ is the initial mass of the 2nd stage,

M _t2 - the mass of fuel of the Central unit spent on the flight section of the 2nd stage,

m _to2 - mass of the fuel compartment of the central unit (2nd stage),

m _dv2 - the mass of the propulsion system of the central unit (2nd stage),

m _CR2 - 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:

where µ _k2 is the relative final mass of the 2nd 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):

where α ₂ is the mass coefficient of the 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):

where γ ₂ is the mass coefficient of the engines of the central block (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:

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):

those.:

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

then

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

Denote:

Then from (27), taking into account (18):

To assess the influence of the average integral throttle coefficient of the engines of the central unit k _{d av} on the relative mass of the payload of the first 2 stages of the launch vehicle µ _mon , we consider from (31) its partial derivative taking into account (16), (28) and (29):

Find the extremum of the function µ _mon = f (k _{d cf} ):

we are interested in the case when:

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

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 α ₂ ≈α ₁ , i.e. (α ₂ / α ₁ ) ≈1 and, therefore, given that all mass coefficients are positive, the right-hand side of expression (34):

At the same time, the left-hand side of expression (34) µ _к2 <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:

This suggests that the partial derivative ∂µ _mon / ∂k _{d cp} <0, i.e. to increase the relative payload mass µ _{bp of the} first 2 stages of the launch vehicle, in which the central and side blocks use the same rocket fuel components, it is necessary to reduce the average integral throttle coefficient of the central unit engines k _{d cf} until they are completely turned off. However, turning off the engines of the central unit is only possible when a steady movement along the trajectory is achieved, therefore, in practice, k _{d cf} will always be greater than zero, otherwise the main idea of the package is lost - the creation of additional starting thrust by the engines of the rocket block of the 2nd stage.

Let us estimate the value of the ratio of the mass coefficients of the fuel compartments of the central and lateral blocks corresponding to the extremum point, after which the partial derivative ∂µ _mon / ∂k _{d av} > 0, and to increase the payload mass µ _mon , on the contrary, it is necessary to increase the average integral throttle coefficient of the engines of the central block k _{d Wed} To do this, we write expression (33) in the form of the inequality:

or:

we write about α ₂ :

or:

and finally:

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

α ₁ = 0.05 ... 0.10,

µ _k2 = 0.3 ... 0.6,

β ₂ = 0.01 ... 0.02.

Then the value of [1 + α ₁ · (1-µ _к2 )] 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 _mon with a decrease in the average integral throttle coefficient of the engines of the central unit k _{d av} , it is necessary that the ratio the mass coefficients of the fuel compartments of the central and side blocks (α ₂ / α ₁ ) 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 (9)

1. A method of putting payload into orbit with a launch vehicle with a multiblock package of missile blocks of a combined scheme, 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 has reached 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 switched 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 the side rocket blocks when the main rocket block rocket engine is 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 ² . 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 mid-range rocket engine of the central missile unit is launched and maintained unchanged 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.

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