RU2386829C1 - Hypersonic turbo ejector engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: hypersonic turbo ejector engine comprises inlet device, compressor, main combustion chamber, turbine gas ejector, afterburner and outlet device. Ejector high-pressure channel communicates, on one side, with compressor via main combustion chamber, and, on its opposite side, it communicates with turbine via mixing chamber. Ejector low-pressure channel communicates, on one side, with atmosphere via main combustion chamber, and, on its opposite side, it communicates with turbine via mixing chamber. Has ejector low-pressure channel accommodates steam-air ejector with its high-pressure channel making a continuation of high-pressure channel of air-to-air heat exchanger arranged at engine compressor outlet.

EFFECT: higher thrust and reduced fuel consumption for entire range of altitudes and speeds.

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Description

The invention relates to aircraft engine manufacturing.

Today, the thrust and flow characteristics of aircraft gas turbine engines (GTE) are inferior to the similar characteristics of ramjet engines, if the flight speed is more than three to five Mach numbers. Moreover, the combination of gas turbine engines and ramjet engines does not improve the characteristics of the original engines.

The aim of the invention is the development of a gas turbine engine with better traction and expense characteristics than ramjet.

Known turbojet engine (Patent RU 2190772, IPC 7 F02C 3/32, 1999) containing an input device, a compressor, a main combustion chamber, a turbine, a gas ejector, the high pressure channel of which is connected on one side to the compressor through the main combustion chamber, and on the other hand, with the turbine through the mixing chamber, the low pressure channel is connected on one side to the atmosphere through the inlet device, and on the other hand, with the turbine through the mixing chamber, afterburner, and outlet device. The disadvantage of turbojet engines (TRDE) is that at flight speeds of more than five Mach numbers the pressure drop (pressure difference at the outlet and inlet) in the turbocharger becomes negative, which makes the traction and flow characteristics of the turbojet engine worse than for ramjet engines.

Heat transfer devices that are used in gas turbine engines are known, including for cooling air at the compressor inlet (Nechaev Yu.N. Power plants of hypersonic and aerospace aircraft. - M.: Cosmonautics Academy, 1996, Fig. 24, 25) .

This goal is achieved by the fact that a steam-air ejector is installed in the low-pressure channel of the gas ejector TRDE, the high-pressure channel of which is a continuation of the high-pressure channel of the water-air heat exchanger installed at the inlet to the engine compressor. The appearance of new elements in the TRE allows to solve a number of technical problems: a) increase the pressure in the low-pressure channel of the gas ejector; b) lower the air temperature at the inlet to the compressor; c) reduce the coefficient of excess air in the main combustion chamber without increasing the gas temperature in front of the turbine, which ultimately allows you to achieve your goal.

Figure 1 shows a diagram of a hypersonic turbojet engine (GTRDE);

figure 2 shows the dependence of the thrust coefficient GTRDE and ramjet from the Mach number;

figure 3 shows the dependence of the overall efficiency of GTRDE and ramjet on the Mach number;

figure 4 shows the dependence of the specific impulse of various jet engines on the Mach number.

GTRDE (figure 1) consists of an input device 1, a steam-air ejector 2, a water-air heat exchanger 3, a low pressure channel 4, a compressor 5, a main combustion chamber 6, a gas ejector with a mixing chamber 7, a turbine 8, an afterburner 9, an output device 10. In this case, the high-pressure channel of the gas ejector is connected to the compressor through the main combustion chamber, and the low-pressure channel to the atmosphere through the inlet 1. The mixing chamber 7 is connected on one side to the gas ejector and, on the other hand, to the turbine 8.

The engine is as follows. Air from the atmosphere through the inlet 1 enters the low pressure channel 4 and into the compressor 5 (through the low pressure channel of the heat exchanger 3). Compressed in the compressor to a predetermined pressure, the air flows in a continuous stream into the main combustion chamber 6, where fuel is simultaneously supplied through the nozzles. The gas generated as a result of combustion enters the high-pressure channel of the gas ejector, ending with a tapering nozzle, and then into the mixing chamber 7. The flow rate increases when the air flows out of the nozzle, and the static pressure drops, which creates conditions for ejection of air from the low-pressure channel 4 into the chamber blending. In the mixing chamber, the air and gas are mixed, braked, as a result of which, at the outlet of the mixing chamber, an increased (relative to the air pressure in the input device) full gas pressure is established. From the mixing chamber 7, the gas enters the turbine 8. The turbine drives the compressor 5. The gas exiting the turbine enters the afterburner, after which it expands in the output device and flows into the atmosphere, creating thrust.

At flight speeds of more than five Mach numbers, water is supplied to the heat exchanger 3 under high pressure (more than 5 MPa) using a hydraulic pump, which evaporates in the heat exchanger channels (water temperature is above critical, which is 648 K for water) and enters the channel in the form of steam high pressure vapor-air ejector 2, from where with sound (supersonic) speed flows into the low pressure channel 4 of the gas ejector. During operation of the heat exchanger 3 (steam-air ejector), energy transformations occur: a) the air temperature in front of the compressor decreases, which facilitates air compression - it increases the gas pressure behind the compressor (in front of the turbine); b) the gas pressure in the low-pressure channel of the gas ejector increases by 15 ÷ 20% (kinetic energy of the vapor is converted to gas pressure), which also increases the gas pressure in front of the turbine; c) the heat capacity of the gas increases, which, at a constant gas temperature in front of the turbine, increases fuel consumption (the coefficient of excess air in the main combustion chamber decreases) - the engine power. The indicated energy transformations make it possible to maintain the pressure drop in the output device of the gas turbine engine higher than in the ramjet, which determines the physical nature of the achieved positive result.

By mathematical modeling performed and to compare the efficacy GTRDE ramjet provided job optimum operating parameters for both engines: GTRDE (pressure ratio π start conditions of _k = 3.6, the ejection rate in the start conditions of m = 0,05; the gas temperature before the turbine Tr * = 2300 K; coefficient of excess air in the main combustion chamber α _ks no more than 1.3; total coefficient of excess air α _Σ = 1.0); Ramjet ramjet (α _Σ = 1,0). The characteristics of the input and output devices are standard for the GLA (Nechaev Yu.N. Power plants of hypersonic and aerospace aircraft. - M .: Academy of Cosmonautics, 1996, Table 3, Table 8.a). The flight path is standard for the GLA (Bulletin of the Academy of Cosmonautics. - M .: Academy of Cosmonautics, 1998, No. 2, p. 153, Fig. 1). The change in the properties of a gas depending on its temperature and composition, as well as the pressure loss in the engine elements and the dissociation of combustion products during the calculation are taken into account.

Figures 2 and 3 show the dependences of the traction coefficient Cp and the overall efficiency on the Mach number for GTRDE and ramjet. It can be seen that in the entire range of flight speeds (up to seven Mach numbers), the traction and discharge characteristics of the engine are better than those of the ramjet. The increase in efficiency at speeds M> 5 (Fig. 3) is due to the use of an additional working fluid in the gas turbine engine - water, the flow rate of which is comparable to the fuel consumption (at the maximum flight speed, water consumption reaches 200% of fuel consumption).

Figure 4 shows the patterns of specific impulses along M for turbojet engines, ramjet engines, engine scramjet engines and liquid propellant rocket engines operating on hydrogen and kerosene, according to TsIAM. Here, for comparison, the specific pulses of TRDE (shaded areas) calculated by the author, including GTRDE, are plotted. The upper boundaries of these regions correspond to the minimum temperatures of the working fluid (maximum efficiency) at which horizontal flight of the aircraft is possible, and the lower ones correspond to the maximum temperatures at which the maximum frontal thrust of the engine is realized. It can be seen that TRDE in the speed range from two and a half to seven Mach numbers are the most efficient engines among the known analogues.

GTRDE can find application in the space industry. A promising direction is the delivery of goods into space using two-stage aerospace systems, in which the first stage is a hypersonic accelerator aircraft, which accelerates the second stage to speeds M = 6 ÷ 7. According to experts, the cost of delivering goods into space when using two-stage aerospace systems is reduced by 5–10 times (Bulletin of the Academy of Cosmonautics. - M: Academy of Cosmonautics, 1998, No. 2, p. 115).

Claims (1)

A hypersonic turbojet engine containing an input device, a compressor, a main combustion chamber, a turbine, a gas ejector, the high pressure channel of which is connected on one side to the compressor through the main combustion chamber, and on the other hand, to the turbine through a mixing chamber, a low pressure channel on one side the side is connected to the atmosphere through the inlet device, and on the other hand, to the turbine through the mixing chamber, the afterburner, the outlet device, characterized in that in the low-pressure gas channel zhektora installed steam-ejector, high pressure channel which is a continuation of the high-pressure channel the water-air heat exchanger mounted at the inlet of the compressor motor.

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