RU2187009C2 - Two-chamber turbojet engine (versions) - Google Patents

Abstract

FIELD: mechanical engineering; engines. SUBSTANCE: turbojet engine has intake device, compressor, turbine, main combustion chamber arranged between compressor and turbine, and ejector nozzle. Ejector nozzle has primary (inner) nozzle connected with compressor by channel accommodating additional combustion chamber, and secondary (outer) nozzle connected with turbine by channel accommodating reheat combustion chamber. Degree of pressure reduction across turbine is 2-2.5 times less than degree of air compression in compressor: 15-20. According to other design version, pressures in main and additional chambers of two-chamber turbojet engine are equal. Outer nozzle is made supersonic. Exit section of inner nozzle is located in widening part of outer nozzle. EFFECT: decreased specific mass of engine. 3 cl, 7 dwg

Description

 The invention relates to aircraft engine manufacturing.

 Known single-circuit turbojet (turbojet) engines (Theory and calculation of the WFD. Edited by S.M. Shlyakhtenko, Moscow, Mechanical Engineering, 1987, p. 16, Fig. 1.1). Engines have an increased specific fuel consumption due to the high temperature of the gas flowing out of the nozzle.

 Double-circuit turbojet engines are known (SU, 312328, 1941). Engines have an increased midship area (due to the presence of a low-pressure fan) and, as a result, an increased specific gravity.

 Forced turbojet engines are known (Theory and Calculation of WFD. Edited by S.M. Shlyakhtenko, Moscow, Mechanical Engineering, 1987, pp. 16-17, Fig. 1.2, 1.4). Engines have reduced (due to reduced pressure in the afterburner) thermal efficiency

 Double-circuit turbojet engines with forcing along the outer contour are known (Yu.N. Nechaev, P.M. Fedorov. Theory of aviation gas turbine engines, part 2, Moscow, Mechanical Engineering, 1978, p. 232, Fig. 17.3 (b)). Engines have a reduced thermal efficiency, subsonic nozzles. In general, engines with forcing along the outer contour are inferior to their counterpart - a dual-circuit turbojet engine with a common mixing chamber (less economical and less speed).

 Known turbojet engine with air intake due to the compressor (Patent GB 1201526, IPC F 02 K 3/02, 1970), which uses an additional combustion chamber located in the turbine space to create additional traction. The disadvantage of the engine is the high specific fuel consumption at cruising engine operating modes, which is a consequence of high heat losses during the outflow of gas into the atmosphere (the temperature of the outgoing gases cannot be lower than the temperature of the gases behind the turbine).

 Sound gas ejectors are known (TN Abramovich. Applied gas dynamics, Moscow, Nauka, 1976, p. 487, Fig. 9.4).

 Famous supersonic gas ejectors (T. N. Abramovich. Applied gas dynamics, Moscow, Nauka, 1976, p. 534, Fig. 9.22).

The closest analogues of the invention are:
1. A turbojet engine with air intake due to the compressor (Patent GB 1201526, IPC F 02 K 3/02, 1970).

 2. Double-circuit turbojet engine with forcing along the outer contour (Yu.N. Nechaev, P.M. Fedorov. Theory of aviation gas turbine engines, part 2, Moscow, Mechanical Engineering, 1978, p. 232, Fig. 17.3 (b)).

 The aim of the invention is to improve the specific (thrust, fuel consumption, mass) engine characteristics.

 This goal is achieved by the fact that the air behind the compressor of the known turbojet engine, in which the degree of pressure reduction on the turbine is 2-2.5 times less than the degree of compression of the air in the compressor, the value of which is 15-20, is divided into two flows, each of which is sent to the chamber combustion (primary and secondary) and further into the ejector nozzle. In this case, part of the energy of one of the flows is spent on mechanical work (compressor drive).

 The essence of the invention lies in the fact that a turbojet engine has two interconnected thermodynamic cycles with heat supply at equal pressures and a common supersonic nozzle in which gas mixing is either absent or occurs on a shortened supersonic section of a common nozzle.

 The presence of two thermodynamic cycles makes it possible to realize to the greatest degree the thermodynamic possibilities of obtaining maximum free operation by a heat engine in the presence of two main restrictions: the gas temperature in front of the turbine blades and the calorific value of the fuel. The interconnection of cycles (changing the ratio of work), provided by quantitative indicators of the degrees of expansion of the gas on the turbine and the compression of air in the compressor, allows you to get a wide range of specific engine parameters. Equal pressures make it possible to use a common compressor and provide a high thermal efficiency. for both cycles. A common nozzle with gas-dynamic flow separation allows for multi-mode operation. and the absence of mixing between streams - minimal loss.

 The invention is implemented by means of schemes (Fig. 1, 2, 3).

Important properties of the circuits are:
1. The possibility of obtaining high specific thrusts (more than 110 kg • s / kg) at relatively low (for forced modes) specific fuel consumption (less than 1.3 kg / kg • h).

 2. The possibility of using a single stage compressor. The optimal degrees of compressor pressure increase (both economic and traction) are: Pk≈15-20. In this case, a change in engine thrust (in a wide range) can be carried out without changing the compressor speed (due to an additional combustion chamber).

 3. The ability to control the compressor by changing the critical section of the nozzle (figure 3).

 4. The possibility of using unregulated internal nozzles (Fig. 1, 2, 3).

 5. The possibility of reducing the mass (midship) of the engine by using the shaft as a channel for supplying air to the additional combustion chamber (Fig. 1).

 6. The possibility of using afterburner (figure 2).

 7. The possibility of applying heat recovery (figure 2). The channels connecting the additional combustion chamber to the compressor are a natural heat exchanger.

 8. The multi-mode supersonic nozzles, which is ensured by gas-dynamic separation of flows (Fig. 1, 2).

 9. The possibility of using mixed fuels with different calorific values (Fig. 1, 2, 3).

 10. Simplicity of design, ease of management.

 Engines (Fig. 1, 2, 3) belong to the category of engines of high thrust for medium flight speeds (Mn less than 2.5).

In FIG. 1 shows a diagram of a turbojet engine;
figure 2 - diagram of a turbojet engine;
figure 3 - diagram of a turbojet engine;
in FIG. 4 - dependence of thrust on flight speed;
figure 5 - dependence of specific fuel consumption on flight speed;
in FIG. 6 - dependence of the thrust coefficient on flight speed;
in FIG. 7 - dependence of total efficiency from flight speed.

 A turbojet engine (Fig. 1) consists of an input device 1, a compressor 2, a main combustion chamber 3, a turbine 4, an additional combustion chamber 5, an ejector nozzle 6. Moreover, the ejector nozzle 6 is supersonic with an adjustable secondary (external) nozzle and an unregulated primary (internal) nozzle. The main combustion chamber 3 is located between the compressor 2 and the turbine 4, and the additional combustion chamber 5 is located inside the channel passing through the engine shaft and connecting the compressor to the entrance to the primary nozzle. The secondary nozzle is connected to the turbine channel.

 The operation of the engine is as follows. Air from the atmosphere through the inlet 1 enters the compressor 2 for compression. Compressed to a given pressure, the air is divided into two streams.

 The first stream enters the main combustion chamber 3, where finely atomized fuel is injected simultaneously through the nozzles. The gas generated as a result of combustion enters the turbine 4, which drives the compressor 2. The gas exiting the turbine enters at a subsonic speed into the secondary nozzle.

 The second stream through the channel inside the shaft enters the additional combustion chamber 5, where at the same time finely atomized fuel is injected through the nozzles. The gas formed as a result of combustion flows at a sound (supersonic) speed from the primary to the secondary nozzle.

 In the subsonic part of the secondary nozzle, the first and second flows move in parallel with acceleration, practically without mixing (G. N. Abramovich. Applied gas dynamics, Moscow, Nauka, 1976, p. 491, Fig. 9.7 (b)). In the critical section, the velocity of the first stream reaches the speed of sound, after which the mixing of the flows begins. At the supersonic nozzle exit, the outflow velocity of both flows reaches its maximum value, which creates a high specific thrust.

 The pressure loss from mixing the flows is minimal, since mixing occurs in a short section (the expanding part of the outer nozzle), and, in practice, it can be assumed that the flows move independently.

где R_{уд трд} - удельная тяга ТРД,
С_{уд трд} - удельный расход топлива ТРД,
Тг₁* - температура газа в основной камере сгорания,
Тг₂* - температура газа в дополнительной камере сгорания.Theoretical studies performed by the author show that the specific parameters of a two-chamber turbojet engine (R _beats - specific thrust. With _beats - specific fuel consumption) can be estimated in relation to the similar specific parameters of a single-chamber turbojet engine by the ratios:

where R _{beats turbojet} - specific thrust turbojet engine,
With _{beats TRD} - specific fuel consumption of the turbojet engine,
Tg ₁ * - gas temperature in the main combustion chamber,
Tg ₂ * is the gas temperature in the additional combustion chamber.

The optimal compression ratio of a two-chamber turbojet engine is defined as:
Port = (0.95-0.97) • Port _TRD ,
where Port _turbojet is the optimum compression ratio of the turbojet engine.

Analysis of the presented ratios shows:
1. The specific thrust and specific fuel consumption in a two-chamber turbojet engine vary, almost proportionally (Tg ₁ * const, Tg ₂ * = var).

раз, где Tг₂*=Tг₂*_max.2. The maximum specific thrust of a two-chamber turbojet engine is higher than the maximum specific thrust of a turbojet engine at approximately

times, where Tg ₂ * = Tg ₂ * _max .

раз, где Tг₂*= Tг₂*_min.3. The minimum specific fuel consumption of a two-chamber turbojet engine is lower than the minimum specific fuel consumption of a turbojet engine

times, where Tg ₂ * = Tg ₂ * _min .

 4. The optimal compression ratio of a two-chamber turbojet engine is 3-5% lower than the optimal compression ratio of a turbojet engine.

 Theoretically, a two-chamber turbojet engine has the best characteristics when the degrees of increase in pressure in the compressor (Pk) and decrease in pressure on the turbine (Fri) are equal. In practice, the best ratio between Pk and Fri is: Pk = (2-2.5) • Fri, which is associated with a limitation in the area of the midsection and ensuring the working conditions of the output device.

Calculations show that with the existing characteristics of the elements of turbojet engines (taking into account losses), a real engine (Pk = 20, Pt = 10, Tg ₁ * = 1600K, α = 1.1) in the starting conditions has the following characteristics:
- when the additional combustion chamber is off: Rд = 68 [dan • s / kg]; Court = 0.851 [kg / dan • h];
- with a working additional combustion chamber: Ryd = 110 [Dan • s / kg]; Cyd = 1.3 [kg / dan • h].

 In flight conditions, for two-chamber turbojet engines, the compressor speed remains constant. The change in traction is carried out by changing the operating mode of the additional combustion chamber. Tapering nozzle (due to high temperatures) - unregulated. In this regard, for engines made according to the schemes (Fig. 1, 2), the turbine must have a working margin of at least 30%, which is necessary to ensure high PCs when working with the auxiliary combustion chamber switched off. High compression ratios (Pk = 15-20) make it possible to use the engine shaft (Fig. 1) as a channel for supplying air to an additional combustion chamber (for an engine with a thrust of 10 tons, the minimum diameter of the channel is 16-20 cm).

 In FIG. 2 shows a diagram of a forced two-chamber turbojet engine. A new engine element is the afterburner 7 of the combustion, located inside the channel connecting the turbine to the secondary nozzle. An additional combustion chamber 5 is connected by a channel, consisting of nozzles 8 and a channel forming an external circuit, with a compressor.

 The operation of the engine is as follows. When a supersonic flight speed (Mp more than 1.5) is reached, finely atomized fuel is injected into the afterburner. As a result of fuel combustion, the following increases: gas temperature, its outflow rate and, accordingly, engine thrust.

 In FIG. 3 shows a diagram of a two-chamber turbojet engine with a supersonic ejector. The engine has an additional combustion chamber 5 located in the outer channel connecting the compressor with an adjustable supersonic (external) nozzle 9, a channel 10 connecting the turbine with an unregulated sound (inner) nozzle 11. The nozzle 11 is located so that its cut is inside the expanding part of the nozzle 9, thereby forming a supersonic gas ejector.

 The operation of the engine is as follows. The air is compressed in the compressor, after which it is divided into two streams. The first stream enters the main combustion chamber and then through the turbine through the channel 10 into the nozzle 11. The second stream enters the additional combustion chamber 5 and further into the nozzle 9. The gas pressure in the first stream (after passing through the turbine) is less than the gas pressure in the second stream.

 The first stream accelerates in the nozzle 11 to the speed of sound (static pressure decreases by the value of the critical differential). The second stream accelerates in the nozzle 9 to a supersonic speed (the static pressure in the cut plane of the nozzle 11 decreases to a pressure lower than in the first stream, which is achieved due to a significantly higher flow rate of the second stream). In the cross section of the nozzle exit 11, flows are encountered, mixed and accelerated.

 The advantage of the circuit of figure 3 is that the operation of the additional combustion chamber does not affect the operation of the compressor. The disadvantage of this scheme is that the gas temperature in the additional combustion chamber is limited due to the need to cool the inner casing. The disadvantage is that it is impossible at the same time (except for the design mode) to ensure full expansion of the gas in both flows, since the pressure drop across the sound nozzle is determined by the operating mode of the supersonic nozzle, which constantly changes under flight conditions. The independence of the nozzle operating modes can be ensured if the sonic nozzle is moved outside the supersonic nozzle, but in this case the sonic nozzle will operate in the mode of maximum underexpansion.

In FIG. 4-7 show the speed characteristic (for an altitude of 11 km) of two-chamber engines made according to the schemes of FIG. 1, 2, 3 with the initial data: thrust at start R = 10 t; Tg ₁ * = 1600 K; Tg ₂ * = 2300 K; TF * = 2000 K; Pk = 20; Fri = 9; the characteristics of the elements are standard. The speed characteristic consists of three characteristic sections: 1 - the main combustion chamber is working; 2 - the primary and secondary combustion chambers work; 3 - the main, secondary and afterburner combustion chambers work (for the circuit of FIG. 2).

 Studies performed by the author show that two-chamber turbojet engines have a sufficiently high gas-dynamic perfection, are simple in design and control. The expected decrease in the specific gravity of the engine (relative to the known turbojet engines) is 25–40%.

Claims (3)

 1. A two-chamber turbojet engine containing an input device, a compressor, a turbine, a main combustion chamber located between the compressor and the turbine and an ejector nozzle, consisting of a primary (internal) nozzle connected to the compressor by a channel inside which an additional combustion chamber is located, and a secondary ( external) nozzle connected to the turbine by a channel, inside which there is an afterburner of combustion, characterized in that the degree of pressure reduction on the turbine is 2-2.5 times less than the degree of compression in zduha in the compressor, which is the value of 15-20.  2. The two-chamber turbojet engine according to claim 1, characterized in that the channel connecting the compressor to the additional combustion chamber passes through the engine shaft.  3. A two-chamber turbojet engine containing an input device, a compressor, a turbine, a main combustion chamber located between the compressor and the turbine, a channel connecting the turbine to the internal nozzle, a channel connecting the compressor to the external nozzle, an additional combustion chamber located in the external channel, characterized in that the gas pressure in the main and secondary chambers is the same, the outer nozzle is supersonic, and the cut of the inner nozzle is located inside the expanding part of the outer nozzle.

Images