RU150686U1 - CARRIER ROCKET FOR LIGHT LOADS - Google Patents

Abstract

1. A booster rocket comprising at least two stages, characterized in that the liquid-propellant rocket engine of the second stage rocket block is based on the combustion chamber of the first stage engine. The launch vehicle according to claim 1, characterized in that the first stage rocket unit consists of a transition compartment, an oxidizer tank, an instrument compartment, a fuel tank and a tail compartment. The launch vehicle according to claim 2, characterized in that the tail compartment of the first stage includes a liquid nitrogen tank, a hydrogen peroxide tank and a marching liquid engine. The launch vehicle according to claim 3, characterized in that the first stage oxidizer tank is charged with gasified oxygen, and the fuel tank with gasified nitrogen. The launch vehicle according to claim 4, characterized in that the oxidant and fuel tanks of the first stage have a smooth cylindrical surface and are made of aluminum alloy AMg6.6. The launch vehicle according to claim 5, characterized in that the RD-108A is used as a liquid-propellant first-stage engine, made in an open circuit with a hydrogen peroxide gas generator. The launch vehicle according to claims 1 to 6, characterized in that the second stage rocket unit consists of an instrument compartment, an oxidizer tank, an inter-tank compartment, a fuel tank and a tail compartment. The carrier rocket according to claim 7, characterized in that the oxidizer and fuel tanks of the second stage are suspended, spherical, from an aluminum alloy AMg6.9. The launch vehicle of claim 8, wherein the inter-tank and tail compartments of the second stage are conical in shape. The launch vehicle of claim 10, characterized in that a march is installed in the tail compartment of the second stage

Description

Technical field

The utility model relates to the field of rocketry and can be used in the development of a launch vehicle (LV) for light loads. State of the art

Over the past 10 years, there has been a tendency to move from heavy satellites weighing several tons to micro and nano-class satellites. The development of micro, mini and nanosatellite platforms are observed around the world. Both private and state-owned companies and educational institutions participate in the creation of devices of such classes. Modern domestic light-class launch vehicles have an excess payload to bring single micro and nanosatellites into orbit, which leads to the need for cluster launches. This practice is inconvenient for customers launching individual devices. Given the fact that modern nanosatellites have a mass of 1-10 kg, and microsatellites from 10-10 kg, there is an urgent need to create an easy and relatively inexpensive launch vehicle for such spacecraft. Currently, many countries and private companies are working in this direction, but not one country or private company does not have a light and relatively inexpensive PH. In the case of using middle-class missiles for the simultaneous launch of micro and nanosatellites, a situation often arises when the delay in creating the main payload leads to the fact that the launch time for associated loads is disrupted. Compliance with the withdrawal schedule is especially critical when deploying orbital groups consisting of several devices. The proposed utility model solves the problem of high efficiency and convenience of launches of spacecraft micro and nano class.

The prior art conversion military three-stage space rocket launcher "Rokot", created on the basis of ballistic missiles UR-100N UTTX. The missile uses highly toxic fuel components - UDMH + AT. There is no unification of the propulsion system at all stages of the rocket. The presence of a special booster block to improve the energy characteristics of a ballistic missile to turn it into a space carrier, which complicates the design and reduces the reliability of the entire carrier. It should also be noted that the ICBM UR-100N UTTX is not currently being produced. For reference, “Rokot” has a commercial value of 1 kg in orbit of 18261 USD.

The prior art conversion military three-stage space launch vehicle "Dnepr", created on the basis of the liquidation of intercontinental ballistic missiles RS-20. The missile uses highly toxic fuel components - UDMH + AT. The presence of a special booster block to improve the energy characteristics of a ballistic missile to turn it into a space carrier complicates the design and reduces the reliability of the entire carrier. There is no unification of the propulsion system at all stages of the rocket. For reference, Dnepr has a commercial cost of 1 kg in orbit of 37500 USD.

The prior art three-stage launch vehicle of the middle class Soyuz-2.1v. The closest analogue of a utility model. The first stage uses the NK-33 rocket engine, which is currently not being produced. The rocket uses a complex structure for a propulsion system. There is no unification of the propulsion system at the rocket stages, which leads to low ease of maintenance and at launch. The use of turbopump units (fuel supply system in a rocket engine) on all engines complicates the design. For reference, Soyuz-2.1v has a commercial value of 1 kg in orbit of 99,000 USD.

The technical problem and the technical result

The technical task is to increase the fault tolerance (reliability) of the launch vehicle blocks, simplifying the design, providing the possibility of mass production and launches. The technical result coincides with the technical task.

Decision

To solve the technical problem, a launch vehicle is proposed that includes at least two stages, while the second-stage liquid-propellant rocket engine is based on the combustion chamber of the first-stage engine. The rocket unit of the first stage may consist of a transition compartment, an oxidizer tank, an instrument compartment, a fuel tank and a tail compartment. The tail compartment of the first stage may include a liquid nitrogen tank, a hydrogen peroxide tank, and a liquid propellant propulsion engine.

The launch vehicle can be designed in such a way that the first stage oxidizer tank is pressurized with gasified oxygen, and the fuel tank with gasified nitrogen. The oxidizer and fuel tanks of the first stage can have a smooth cylindrical surface and be made of aluminum alloy AMg6.

The launch vehicle can be designed in such a way that RD-108A is used as a liquid-propellant jet engine of the first stage, made according to an open circuit with a gas generator based on hydrogen peroxide. The second stage unit may consist of an instrument compartment, an oxidizer tank, an inter-tank compartment, a fuel tank and a tail compartment. In this case, the oxidizer and fuel tanks of the second stage can be suspended, spherical, from the aluminum alloy AMg6.

The launch vehicle may be designed such that the inter-tank and tail section of the second stage are conical in shape. In the tail compartment of the second stage, a marching liquid propellant engine based on the RD-108A with fuel displacement and carbon-fiber nozzle nozzle can be installed.

Description of drawings

The essence of the utility model is illustrated in Fig., Which shows the structural and layout diagram of the launch vehicle "Adler". The following notation is introduced

1 - head fairing,

2 - payload,

3 - instrument compartment

4 - oxidizer tank,

5 - inter-tank compartment,

6 - fuel tank,

7 - tail compartment

8 - LRE based on the RD-108 camera,

9 - transition compartment

10 - oxidizer tank,

11 - instrument compartment

12 - fuel tank

13 - tank of liquid nitrogen,

14 - tank of hydrogen peroxide,

15 - RD-108A LPRE,

16 - tail compartment.

Detailed Solution Description

The solution is a launch vehicle comprising at least two stages, characterized in that the liquid-propellant rocket engine of the second-stage rocket unit is based on the combustion chamber of the first-stage engine. In this case, the rocket unit of the first stage consists of a transition compartment 9, an oxidizer tank 4, an instrument compartment 3, a fuel tank 6 and a tail compartment 7, which includes a liquid nitrogen tank, a hydrogen peroxide tank and a marching liquid propellant engine (see Fig.). The oxidizer tank of the first stage is pressurized with gasified oxygen, the fuel tank with gasified nitrogen. To simplify the technological costs, the oxidizer and fuel tanks of the first stage have a smooth cylindrical surface and are made of aluminum alloy AMg6. As a liquid-propellant jet engine of the first stage, RD-108A is used, made according to an open circuit with a gas generator based on hydrogen peroxide. This liquid-propellant rocket engine with four steering chambers uses environmentally friendly components, oxygen and kerosene as fuel, and has been mass-produced since 1957, demonstrating high reliability. The use of a serial, mastered in production, rocket engine provides a high degree of reliability (reliability) and ensures high adaptability of utility model rocket engines.

The rocket block of the second stage consists of the instrument compartment 11, the oxidizer tank 10, the inter-tank compartment, the fuel tank 12 and the tail compartment 16. The oxidizer and fuel tanks of the second stage are suspended, spherical, made of aluminum alloy AMg6, and the inter-tank and tail compartment of the second stage are conical in shape . At the same time, in the tail compartment of the second stage there is a marching liquid propellant engine created on the basis of RD-108A with a displacement fuel supply and a carbon-plastic nozzle nozzle.

The creation of a two-stage launch vehicle using a first-stage engine combustion chamber for second-stage engines can significantly simplify and reduce the number of elements, eliminating additional booster blocks and stages with a slight increase in fuel mass (for reference, the cost of fuel is about 4-5% of the cost of the entire rocket -carrier), and provide the possibility of mass production and launches.

Implementation example

As an example of implementation, the Adler launcher, which consists of two missile units connected in a tandem scheme, is shown. At all stages, liquid propellant rocket engines are used on environmentally friendly fuel components: oxidizing agent - liquid oxygen, fuel - kerosene. The rocket unit of the first stage consists of a transition compartment, an oxidizer tank, an instrument compartment, a fuel tank and a tail compartment. The oxidizer and fuel tanks are cylindrical, smooth, made of aluminum alloy AMg6. The tail compartment includes a liquid nitrogen tank, a hydrogen peroxide tank and a marching rocket engine. The RD-108A liquid propellant rocket engine (used on the central unit of the Soyuz launch vehicle) is designed according to an open circuit with a hydrogen peroxide gas generator. The oxidizer tank is pressurized with gasified oxygen, the fuel tank with gasified nitrogen. The control at the stage of flight of the first stage is carried out due to the rotary steering chambers of the marching rocket engine in all three axes. The missile unit of the second stage consists of an instrument compartment, an oxidizer tank, an inter-tank compartment, a fuel tank and a tail compartment. The oxidizer and fuel tanks are suspended, spherical, made of aluminum alloy AMg6. The inter-tank and tail section are conical. In the tail compartment there is a marching liquid propellant rocket engine created on the basis of a single combustion chamber of the RD-108A serial engine with a displacing fuel supply and a carbon-plastic nozzle nozzle. Control over the channels of the course and pitch is carried out by installing a marching rocket engine in a universal joint suspension, along the roll - due to two rotary control engines. The separation of the steps occurs according to the cold scheme using explosive pyro-bolts. The table shows the main parameters and characteristics of the rocket.

Claims (10)

1. The launch vehicle, comprising at least two stages, characterized in that the liquid-propellant rocket engine of the rocket block of the second stage is based on the combustion chamber of the first stage engine. 2. The launch vehicle according to claim 1, characterized in that the rocket unit of the first stage consists of a transition compartment, an oxidizer tank, an instrument compartment, a fuel tank and a tail compartment. 3. The launch vehicle according to claim 2, characterized in that the tail compartment of the first stage includes a tank of liquid nitrogen, a tank of hydrogen peroxide and a marching liquid engine. 4. The launch vehicle according to claim 3, characterized in that the first stage oxidizer tank is charged with gasified oxygen, and the fuel tank with gasified nitrogen. 5. The launch vehicle according to claim 4, characterized in that the oxidizer and fuel tanks of the first stage have a smooth cylindrical surface and are made of aluminum alloy AMg6. 6. The launch vehicle according to claim 5, characterized in that the RD-108A, made in an open circuit with a gas generator based on hydrogen peroxide, is used as the first-stage liquid-propellant engine. 7. The launch vehicle according to claims 1 to 6, characterized in that the second stage rocket unit consists of an instrument compartment, an oxidizer tank, an inter-tank compartment, a fuel tank and a tail compartment. 8. The launch vehicle according to claim 7, characterized in that the oxidizer and fuel tanks of the second stage are suspended, spherical, from an aluminum alloy AMg6. 9. The launch vehicle of claim 8, wherein the inter-tank and tail compartments of the second stage are conical in shape. 10. The launch vehicle of claim 10, characterized in that in the tail compartment of the second stage there is a marching liquid propellant engine created on the basis of RD-108A with a fuel displacement and carbon-fiber nozzle nozzle.

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