RU2799263C1 - Integrated direct-flow air-jet engine - Google Patents

Abstract

FIELD: aviation; engines.

SUBSTANCE: proposed device relates to jet engines, in particular to integral (combined) ramjet engines (ramjet), and is designed to equip disposable and reusable aircraft. The task to be solved by the proposed device is to increase the reliability and efficiency of the propulsion system, to ensure the possibility of multiple use of aircraft with a ramjet. The integrated ramjet engine includes a main ramjet engine and a starting liquid propellant engine (LPE) located in its combustion chamber. The section of the LPE nozzle exit is in the plane of the critical section of the propulsion ramjet. The nozzle exit area of the launch rocket engine is less than the critical section area of the propulsion ramjet engine by the size of the air intake throat area. In the acceleration mode of the aircraft, the launch rocket engine operates. The air passing through the ramjet chamber passes through the annular gap formed by the LPE nozzle exit and the critical section of the ramjet. In the expanding part of the ramjet nozzle, a nozzle with a break in the supersonic contour is formed. After the end of the acceleration of the aircraft to the cruising speed of the ramjet, the LPE is moved to the front of the ramjet chamber, it is turned off and fixed. The LPE fuel supply system can be used to spray it into the ramjet combustion chamber in addition to the ramjet's own fuel supply system.

EFFECT: reusable engine, higher specific impulse of the accelerating engine compared to solid propellant rocket motors, decrease in the aerodynamic resistance of the air intake device in the accelerating section of the flight, and an increase in reliability due to the absence of destructible nodes.

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Description

The proposed device relates to jet engines, in particular to integral (combined) ramjet engines (ramjet engines), and can be used as a propulsion system for aircraft.

Currently, ramjet engines (PRjet) are used in aircraft. A significant disadvantage of such engines is the inability to work at the start and at a relatively low flight speed (usually M<2). To quickly achieve the value of the speed of the aircraft, required for the effective operation of the sustainer engine (ramjet), a rocket engine is usually used (starting accelerating engine - usually solid propellant rocket engine). Due to severe restrictions on the overall weight characteristics of aircraft with ramjet, the booster rocket engine is located in the ramjet chamber. Such engines are called integral (IPVRD) or combined.

The advantage of the ramjet engine over the RPD (ramjet-ramjet engine), which has intermediate characteristics between rocket and ramjet engines, is higher traction characteristics in the acceleration section and higher efficiency in the cruising section at a flight speed of M≥2.5 [1].

The optimal geometrical characteristics of the nozzle of the sustainer and starting accelerating engines differ significantly. The gas-dynamic operating conditions of the accelerating engine in the launch section and the ramjet in the main section require different chamber nozzle geometry - the critical section area of the ramjet nozzle must be much larger than the critical section area of the starting engine nozzle. The marching nozzle is unsuitable for use in the acceleration mode due to overdimension.

One way to solve this problem is to use a burnable nozzle. A nozzle device for a rocket with a ramjet engine [2] is known, where the starting nozzle burns out under the influence of high-temperature fuel combustion products in the cruising mode of the ramjet engine. The resistance of the starting nozzle in the acceleration mode is ensured by cooling its fire wall due to the supply of coolant through special channels of the firing wall of the starting nozzle in the acceleration mode. At the end of the operation of the starting engine, the refrigerant supply stops. The refrigerant heated in the cooling path is discharged overboard the aircraft. The disadvantage of the system is the presence of a consumable refrigerant and its supply system - this significantly worsens the overall weight characteristics of the aircraft

In [3], a part of the solid fuel charge with a relatively low burning rate is used as a burnable nozzle. A significant nozzle peak occurs from the very beginning of the operation of the accelerating engine, the fuel burn-up characteristics also significantly depend on the fuel charge temperature, which can vary over a wide range. This increases the instability of the overclocking characteristics. The working process is essentially non-stationary and is difficult to predict. This leads to a significant decrease in the thermodynamic efficiency of the engine.

To a lesser extent, the disadvantages of engines with a burnable nozzle are inherent in the device claimed in [4], in which the burnable nozzle of a combined rocket-ramjet engine, located in the internal cavity of the cruise mode nozzle, is made of at least two elements connected to each other with the possibility of forming the path of the booster nozzle from subsonic to transonic and from transonic to supersonic regions.

A common disadvantage of all these devices is an increase in the area of the critical section of the accelerating engine during acceleration, which leads to a decrease in the specific impulse.

In the patent [5], after the operation of the accelerating engine, the plugs of the engine air intake are opened and, under the pressure of the incoming air flow, the idle accelerating engine is removed from the combustion chamber. At the same time, the level of strength of fasteners should ensure reliable fastening of the accelerating engine in the cruising mode and at the same time be destroyed under the action of air flow. Simultaneously with the removal of the accelerating engine, the coaxial outer shell is extended, fastened to the main engine nozzle. This ensures the required length of the once-through combustion chamber. The main disadvantage of this device is the presence of an outer coaxial shell on the ramjet body, which must meet the requirements for ramjet chambers in terms of strength and heat resistance. With large ramjet sizes, this will lead to a significant increase in its weight.

As the closest analogue, the integrated ramjet of the 3M9 anti-aircraft guided missile, designed for use as part of the 2K12 air defense system, was adopted. Propulsion rocket engine - ramjet.

An accelerating engine was located in the ramjet combustion chamber - a solid propellant charge with armored ends, equipped with a Laval nozzle. The accelerating engine nozzle is located inside the propulsion engine nozzle. The solid propellant charge of the accelerating engine is mounted in the propulsion engine chamber using a fiberglass grate. Since the operation of the starting engine is possible only in a closed volume - the exit of combustion products should only be towards the nozzle in the starting section, before the main engine is turned on, the outlets of the air intake channels into the afterburner were closed with fiberglass plugs. Upon completion of the work of the starting engine, it was planned to shoot the inside of the nozzle apparatus (the nozzle of the accelerating engine) with a fiberglass grate holding the starting charge. [6].

The P-100 (3M80) rocket engine of the Moskit anti-ship missile system is arranged in a similar way. The initial acceleration of the Mosquito rocket to the speed necessary for the operation of the ramjet engine was provided by a solid-fuel booster engine located inside the ramjet combustion chamber. In 3-4 seconds after the start, the powder engine burns out and is pushed out of the ramjet nozzle by the oncoming air flow [7].

Recently, one of the requirements for a number of developed HPJEs is the absence of drop elements [4]. The accelerating engine nozzle, as a rule, has a significant mass and its unpredictable movement in the environment is undesirable.

The disadvantage is also the presence of plugs on the air intake directly in front of the combustion chamber of the engine, which leads to an increase in aerodynamic resistance in the starting mode due to the presence of an open path of the air intake device. This air intake device has a complex design of the plug and its removal mechanism and requires a large amount of development and testing, while there remains the possibility of fragments of the plug getting into the engine combustion chamber, which can lead to damage or destruction of the control and fuel distribution elements of the propulsion engine and, ultimately , to the inoperability or destruction of the engine itself. So, for example, the MBDA Meteor air combat missile, during tests after launch, could not switch to the march mode and collapsed in flight. When studying the wreckage, it was concluded that the engine air intakes did not open [8].

It should be noted that in all the analogues considered above, solid propellant rocket engines are used as a starting engine, which have a lower specific impulse compared to LRE and do not have the ability to change thrust in flight, unlike LRE, and are practically “disposable”.

The objective of the proposed technical solution is to create an integrated ramjet engine with increased energy efficiency and the possibility of using both disposable and reusable aircraft.

An integral ramjet engine is proposed, which includes a sustainer ramjet engine and a starting rocket engine located in the ramjet combustion chamber coaxially with it. The nozzle section of the launch rocket engine coincides (is located) in the plane of the critical section of the sustainer ramjet. The area of the nozzle cut of the starting engine is less than the area of the critical section of the sustainer ramjet approximately by the area of the throat of the air intake device. LRE has the ability to move along the axis of the engine to the front of the IVRRE. There are no plugs available in the closest analogue, because fuel combustion in the starting mode does not occur in the ramjet chamber, but in the LPRE chamber. The main elements of the proposed device are shown in Fig. 1 and FIG. 2.

In FIG. 1 shows the ramjet at the starting leg of the flight.

In FIG. 2 shows the ramjet on the cruising leg of the flight.

- air intake device - 1;

- ramjet chamber body - 2:

- fuel injectors of the direct-flow chamber - 3;

- accelerating rocket engine - 4;

- ramjet nozzle - 5;

- LRE-6 movement mechanism.

The movement mechanism includes a telescopic hydraulic cylinder for the extension of the rocket engine and telescopic guides. Cleaning LRE in the front part at the end of the starting section is carried out due to the thrust force of the LRE.

Well-known elements are not shown: the fuel lines of the accelerating and sustainer engines, combustion stabilizers, igniters, as well as devices for fixing the liquid-propellant rocket engine at the starting and sustaining sections of the flight.

Device operation.

At the launch site, the IPVRD works as follows. In the acceleration mode, the starting rocket engine works. The air passing through the air intake and the ramjet chamber passes through the annular gap formed by the LPRE nozzle exit and the critical section of the ramjet. In this case, in the expanding part of the ramjet nozzle, a nozzle is formed with a break in the supersonic contour or the so-called. ejector nozzle [9]. Thus, the LRE combustion products expand not only in the LRE nozzle, but also in the ramjet nozzle, which makes it possible to increase the specific impulse of the starting engine, compared with a shortened nozzle.

After the end of the acceleration of the aircraft, the rocket engine moves to the front of the ramjet chamber. The supply of oxidizer to the LRE chamber is stopped. If necessary, fuel is supplied to the ramjet combustion chamber from the nozzle head of the rocket engine, which, having passed through the nozzle, mixes with air. At the same time, fuel is supplied to the fuel injectors located in the ramjet chamber. The engine is running in cruising mode.

If it is necessary to reduce the flight speed to such a value when the operation of the sustainer engine is impossible, the LRE is moved towards the nozzle of the ramjet engine and the LRE is launched - the aircraft functions as a rocket plane.

Despite the outward similarity of the proposed engine with the RPD, it is not such, because. there is no process of mixing LRE gases with the air of the direct-flow circuit. The placement of the liquid-propellant rocket engine nozzle at the outlet of the direct-flow circuit makes it possible to fully realize the advantages of the liquid-propellant rocket engine (high specific impulse and the ability to control) in the acceleration section.

The use of the proposed IPVRD allows:

- to improve the flight performance of the aircraft due to the possibility of regulating and switching off and on the thrust of the accelerating engine, including ensuring reusable use;

- increase the reliability of the aircraft due to the absence of destructible units (nozzles, air intake plugs and grilles);

- reduce the frontal resistance of the air intake device in the acceleration section.

A possible increase in the mass of the aircraft due to a non-resettable liquid-propellant rocket engine is compensated by a decrease in the mass of the direct-flow chamber, since its strength is calculated based on the operating conditions of the ramjet engine in the cruising section of the flight, and not in the starting mode, in which the ramjet chamber is actually a rocket engine chamber, as well as more high specific impulse of the starting engine.

An assessment of the layout possibility of the proposed IVRJD is presented below.

A ramjet is taken as a main engine, the calculation of which is given in the book by M.M. Bondaryuk [10] pp. 320-324. The parameters of this engine are: thrust - about 2 tons, at a flight speed corresponding to the Mach number 4, at an altitude of 25 km, the cross-sectional area of the neck of the air intake device is 0.28 m ² , the throat area of the nozzle is 0.48 m ² , the nozzle cut area is 1 m ² . LRE is integrated into this engine. The thrust of the accelerating engine is taken from the generally accepted values for ramjet cruise missiles. So, for example, the Burya strategic cruise missile had liquid-propellant rocket engines as launch boosters, providing a starting thrust of about 136 tons. Marching ramjet RD-012 designed by OKB-670 M.M. Bondaryuk provided thrust of 7.75 tons (according to some sources, about 13 tons) [11]. That is, the thrust of the accelerating engine exceeds the thrust of the ramjet in the cruising flight mode by 10 ... 18 times. The same proportions have similar cruise missiles of a potential enemy (Navajo, etc.). Proceeding from this, the thrust value of the booster LRE is assumed to be about 30 tons.

The LRE parameters are determined approximately from the expression [12].

R=Kp×rk×Fcr., where R is thrust, Newton; pk is the pressure in the LRE combustion chamber, Pa; FKp - area of the critical nozzle of the LRE, m ² ,

Kp is the coefficient of increase in the jet force by the nozzle, depends on the polytropic index of the combustion products and the degree of expansion of the nozzle. With a moderate degree of expansion of the nozzle (no more than 100), the average value of Kp is 1.7 [12] (Fig. 2.7).

To obtain such thrust, the area of the critical nozzle of the LRE is required to be about 0.015 m ² (150 cm ² ). According to the layout conditions, the cut-off area of the liquid-propellant rocket engine nozzle should be less than the critical nozzle area of the ramjet engine by the cross-sectional area of the air intake throat (0.28 m ² ), which will be 0.20 m ² . Thus, the expansion ratio of the LRE nozzle is 0.20/0.015≈13. Using the tables of gas-dynamic functions [1], we obtain the ratio of the pressure in the chamber to the pressure at the nozzle exit of about 160, which provides a quite acceptable specific impulse of the LRE.

Information sources

1. Zuev B.C., Macaron B.C. Theory of direct-flow and rocket-ramjet engines, M., "Engineering", 1971, 308 pages.

2. Japan Patent No. 3143654, IPC F02K 7/18, published 04/15/1993

3. US patent No. 4574700, IPC C06D 5/06, publ. 1984

4. RU 2507409 C1 Russian patent 2014 according to IPC F02K 7/18 F02K 1/82.

5. RU 2621588 C1 (IPC F02K 7/18. 2016-07-10.

6. Anti-aircraft missile system 2K12 Kybmissilery.info.

7.A.V. Karpenko. http://bastion-opk.ru/moskit-p-100/ OVT "WEAPON OF THE FATHERLAND".

8. Nevsky Bastion, nevskii-bastion.ru.meteor-mbda.

9. V.D. Sokolov. Momentum loss in jet nozzles with a break in the supersonic circuit. Journal Uchenye zapiski TsAGI 1971, No. 6, star 98-101.

10. Bondaryuk M.M., Ilyashenko S.M. Direct-flow air-jet engines. Publisher: Defense Industry: 1958. 394 p.

11. Product "350" "Storm" strategic cruise missile / project history. www.airbase.ru/sb/russia/lavochkin/la/350.

12. Volkov E.B., Golovkov L.G., Syritsyn T.A. Liquid rocket engines. M., Military Publishing, 1970, 592 p.

Claims (2)

1. Integral ramjet engine, including a sustainer ramjet engine and an accelerating rocket engine installed in the combustion chamber, and the accelerating engine nozzle removed after reaching the cruising speed is integrated into the ramjet nozzle, the critical section areas of these engines are selected from the condition of the engine operation mode close to to the calculated ones, characterized in that an LRE is used as an accelerating engine, equipped with a device that ensures its movement after acceleration to the ramjet cruising speed to the front of the combustion chamber or back to its original operating position for subsequent launch of the LRE. 2. Integral ramjet engine according to claim 1, characterized in that the section of the liquid-propellant rocket engine nozzle is in the plane of the critical section of the sustainer ramjet, and the area of the nozzle exit of the starting rocket engine is less than the area of the critical section of the sustainer ramjet by the size of the throat area of the air intake device.

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