RU2211938C1 - Method of operation of liquid-propellant rocket engine with closed steam-liquid contour in turbopump delivery system - Google Patents

Abstract

FIELD: rocket engines. SUBSTANCE: according to proposed method working steam for operation of turbine is obtained from water using heat of rocket propellant combustion products. Steam is operated to pressure of 0.6-4.0 MPa. Used steam is cooled using cold of oxidizer and condensate is returned into water pump. EFFECT: prevention of freezing of water condensate at high pressure level in chamber. 2 dwg

Description

 The invention relates to liquid-propellant rocket engines (LRE), specifically to a turbo-pumped liquid propellant rocket engine, consisting of separately stored oxidizer and fuel; at least one of these fuel components (oxidizing agent) is cryogenic.

 A known method of operation of a liquid-propellant rocket engine in a turbopump supply system, in which working steam for a turbine is obtained from water using heat from the products of rocket fuel combustion, and the generated steam is cooled using an oxidizer coolant, and the condensate is returned to the water pump: see patent RU 2155273 C1, 08/18/1999, figure 1 - prototype of the invention.

The method of operation of a closed-liquid-liquid propellant liquid propellant rocket engine in a turbopump supply system includes a Rankine duty cycle, at the beginning of which superheated steam is obtained to perform the necessary work on the turbine, and at the end of the cycle, the unused heat of the spent steam entering the liquid propellant rocket is cooled to condense (condense ) it to the extent guaranteeing cavitation-free operation of the corresponding pump. Ultimately, they provide a heat and energy balance between several functional elements (systems) of the liquid propellant rocket engine:
- sources (system of generation) of heat to obtain the working steam of the turbine;
- turbopump unit (TNA) for supplying working fluid;
- a cold source (cooling system) for condensation of spent turbine steam.

 The functioning of each of these elements requires, in turn, the observance of an internal balance between the incoming and outgoing energy. In particular, in TNA, it is necessary to ensure equality between the available capacity of the turbine and the total capacity of the oxidizer pumps, fuel, and condensed product (in our case, water).

The principal advantage of the operation of a liquid-propellant liquid-propellant liquid propellant rocket engine in a turbopump supply system is the absence of loss of specific thrust impulse (I _у ) of the engine to the TNA drive, since exhausted (i.e., energy-depreciating) steam turbines are fed back to the working circuit after condensation, rather than thrown out of the engine through the exhaust system. Due to this, the values of I _y for the engine and for the camera (synonymous with the concept of "traction chamber") coincide. This advantage of the prototype method is combined with moderate pressure of fuel pumps characteristic of it, which makes the development and manufacture of liquid propellant engines cheaper.

. Этот малый расход вместе с высокой плотностью водяного конденсата обусловливаeт малые затраты энергии на привод насоса для рабочего тела турбины (свыше 90% мощности которой может расходоваться в полезных целях - на привод топливных насосов). В итоге энергетический баланс ЖРД может достигаться при высоком уровне давления в камере (р_к), которому соответствуют высокие значения параметра I_у. Малая величина

облегчает также задачу конденсации пара.When choosing a water vapor turbine for a working fluid, it is assumed that water can be heated without fear of dissociation up to a temperature at which the turbine design remains intact. Uncooled gas turbines used in modern rocket engines operate at an inlet temperature of the working fluid T _from = 600 - 1100 K. Water vapor heated to such a high temperature gives a large amount of adiabatic operation on the turbine (L _ad ), and therefore, high power is provided turbines with low water consumption

. This low flow rate, together with the high density of water condensate, results in low energy costs for pump drive for the turbine working fluid (over 90% of the power of which can be spent for useful purposes - for fuel pump drive). As a result, the energy balance of the rocket engine can be achieved at a high level of pressure in the chamber (p _to ), which corresponds to high values of the parameter I _y . Small value

also facilitates the task of condensation of steam.

However, with all of the advantages of the prototype method using an aqueous vapor-liquid circuit, when using this method in practice, it turns out to be impossible to fully use the potential chemical energy of rocket fuel to obtain high values of I _y . This disadvantage is predetermined by the fact that when implementing the known method there is a danger of local freezing of water condensate in the process of its formation at the final stage of the Rankine working cycle. This danger arises from the very low temperature of the liquid oxygen oxidizing agent, which condenses the steam exhausted on the turbine.

It is in view of this danger that the prototype method preferred ammonia over water, despite its inherent restriction on heating (the equilibrium temperature of ammonia dissociation is only 520 K: see the mentioned patent RU 2155273 C1, 08/18/1999, columns 10, 11). At the same time, it was proposed to prevent freezing of condensate to preheat the cryogenic oxygen cooler with the combustion products of rocket fuel. At the same time, however, the remaining refrigeration resource is enough to condense only a small flow rate of the exhaust steam, and due to such a restriction on the flow rate of the turbine working fluid, the energy balance of the LRE is ensured only at a low pressure in the chamber (less than 10 MPa). This leads to low values of the parameter I _y for the prototype method, which deprives it of practical expediency.

 The invention solves the technical problem of preventing freezing of water condensate at a high pressure level in the chamber.

 The stated technical problem is solved in that in the operation method of a liquid-propellant liquid-propellant rocket engine in a turbopump supply system, in which working steam for a turbine is obtained from water using heat from the products of rocket fuel combustion, and the generated steam is cooled using an oxidizer coolant, and the condensate obtained is returned to the water pump, according to the invention, the steam is triggered to a pressure of 0.6-4.0 MPa.

 When carrying out the invention, a technical result is expected that coincides with the essence of the problem being solved.

The invention is illustrated in figures 1 and 2, which presents:
- in FIG. 1 is a functional diagram of a rocket engine made in accordance with the proposed method;
- figure 2 - graphs of the relationship for a number of characteristics of the water vapor-liquid circuit.

 According to figure 1, the liquid propellant rocket engine produces a traction force chamber 1 with a nozzle head 1A, a combustion chamber 1B, and a supersonic jet nozzle 1B. The camera body is formed by two coaxial shells (external and internal), forming a 1G path for the cooler duct. To supply fuel in the engine, a TNA is provided, which contains a two-stage pump of an oxygen oxidizer (liquefied oxygen) 2, a two-stage pump of hydrocarbon fuel (for example, kerosene) 3, a water pump 4 and a steam turbine 5. In order to obtain a working steam of the turbine, a gas generator 6 with a mounted at its exit, heat exchanger-heater 7; they are communicated by the gas duct 8 with the nozzle head 1A of the chamber. Thus, the gas generator, the heat exchanger-heater and the chamber are parts of a common gas-dynamic channel. The nozzle head 6A of the gas generator is connected to the second stage of the pump 2 by means of a pipe 9 and is also connected to the second stage of the pump 3 via the line 10 with the throttle 11. The nozzle head 1A of the chamber is connected to the first stage of the pump 3 through the line 12 with the throttle 13 and is also connected to the first stage of the pump 2 through the line 14 with the heat exchanger-condenser 15 located in it. The line 14 can communicate (after the heat exchanger-condenser) by means of a pipe 14a with otrennym in gazovode 8 ballasting belt 8a.

 To drive the TNA, a closed vapor-liquid circuit of the water working fluid is provided, including the aforementioned pump 4 and a turbine 5. The output of the water pump is connected to the input of the heat exchanger-heater 7 by two parallel lines. One of them includes a consecutively located pipeline 16, a cooling channel 1G of the chamber and a pipe 17. Another highway includes a pipe 18, a cooling jacket 6B of the gas path 6-7-8, a pipe 19, an heat exchanger-economizer 20, and a pipe 21. Its output together with the output the pipe 17 is connected to the mixer 22, and that through the pipe 23 is communicated with the input of the heat exchanger-heater 7. The output of the latter is connected by the pipe 24 to the input of the turbine 5. Its output is communicated by the pipe 25 with the input of the heat exchanger-economizer 20, the output otorogo connected by a pipe 26 to the input of the heat exchanger-condenser 15. Its output is communicated by a pipe 27 to the input of the pump 4. Thus, the pump 4 together with the turbine 5, heat exchangers 7, 15, 20, the cooling path 1G of the chamber and the corresponding connecting lines forms a closed loop for circulation of a water working fluid undergoing phase transformations. In order to regulate the power of the turbine in the circuit provides a bypass steam line with a throttle 28.

 The described LRE works as follows. The rocket fuel oxidizer (liquefied oxygen) enters the pump 2, from which part of the liquid is fed through the pipe 9 to the nozzle head 6A of the gas generator 6. There along the line 10, part of the incoming (kerosene) fuel is fed from the pump 3, which burns in the oxidizer with an excess of the latter . The generated oxidizing gas (with an initial temperature of about 2000 K) enters the path of the heat exchanger-heater 7 and then through the gas duct 8 to the nozzle head 1A of the chamber. The rest of the mass of fuel consumed by the liquid propellant rocket engine is fed there along lines 12 and 14 from pumps 3 and 2. This mass is burned in the combustion chamber 1B with oxidizing gas, and the resulting high-temperature products enter the jet nozzle 1B, flowing out of which, they create a draft. If necessary, part of the oxidizer from the pump 2 after passing through the heat exchanger-condenser 15 can be introduced through the pipe 14a through the ballasting belt 8a into the gas duct 8, thereby reducing the temperature of the oxidizing gas before it is fed to the nozzle head 1A.

 The water working fluid circulating in a closed circuit is distributed after pump 4 along two parallel lines. A part of the total water flow rate is supplied through pipeline 16 to the cooling path 1G of the chamber, where when heated, the liquid is turned into steam, which enters through the pipe 17 to the mixer 22. The rest of the water working medium pumped by the pump 4 also goes there. This part is pumped by the pipeline 18 to the cooling jacket 6B of the gas path 6-7-8 and then through the pipe 19 to the heat exchanger-economizer 20. Here, the incoming water is additionally heated by the exhaust steam of the turbine 5 (see below), after which it is sent to the mixer 22 via the pipe 21 From it, the total mass of the water working fluid enters through the pipe 23 to the path of the heat exchanger-heater 7. The superheated steam obtained in it is fed through the pipe 24 to the turbine 5, leading the pumps 2, 3, 4. The steam is activated on the turbine to a pressure of 0.6- 4.0 MPa, after which it is fed through a pipe 25 to a heat exchanger-economizer 20. Here, the spent steam gives part of its heat to a fresh working fluid, and then the steam is fed through a pipe 26 to a heat exchanger-condenser 15. Here, due to heat exchange with a liquid oxygen oxidizer, steam turned into a liquid, and The resulting water condensate is discharged through the pipeline 27 to the inlet of the pump 4. Next, the described cycle of the water working fluid is repeated.

 The regulation of the LRE is carried out by acting on the throttles 11, 13 and 28. The first of them provides the temperature setting of the gas generator 6. By acting on the second throttle, synchronous emptying of the fuel tanks in the aircraft with the LRE is achieved. The third throttle is used to adjust and regulate traction.

The invention is not limited to the scheme shown in figure 1; in specific cases, other private technical solutions may be used:
- the exhaust steam of the turbine can be cooled not only by an oxidizing agent, but also by combustible rocket fuel;
- the number of impellers in the pumps and turbine may be different;
- a heat exchanger-economizer (20) may not be needed;
- optional is a heat source additional to the chamber for heating the turbine working fluid; figure 1, this additional source is made in the form of an oxidizing gas generator (6); in the case of generating a reducing gas, it can be diluted in the gas duct (8) with liquid fuel, etc.

The problem of freezing water condensate is solved in the invention by optimizing the used Rankine cycle in terms of the pressure at the turbine outlet (p _2T ). In FIG. Figure 2 shows how this parameter affects the following important characteristics:
- the relative flow rate of the aqueous working fluid in the vapor-liquid circuit (m _rel );
- the relative working surface of the heat exchanger-condenser (s _rel ) countercurrent circuit;
- the minimum wall temperature of the heat exchanger-condenser from the side of the water working fluid (t _article ).

 Before commenting on the presented graphs, we recall that ensuring the energy balance of the supply system is only one of the conditions for the operation of a liquid-propellant rocket engine with a closed vapor-liquid circuit. Another necessary condition is the sufficiency and practical feasibility of using the fuel coolant to condense the exhaust steam of the turbine with the return of the resulting water condensate to the operating cycle. In the general case, the main coolant of the fuel is concentrated in a cryogenic oxygen oxidizer having a temperature of ≈100 K at the pump outlet. The first critical factor in the technical implementation of the method under consideration is a heat exchanger-condenser, the dimensions and mass of which can reach unacceptable values. This unit generally contains three working sections: cooling the superheated water vapor to the saturation temperature or the onset of condensation (section 1), condensation itself (section 2) and cooling the condensate to ensure cavitation-free operation of the water pump (section 3).

, и температура отработавшего пара (Т_2T) снижается, что выгодно для габаритно-массовых характеристик теплообменника-конденсатора; но соответственно снижается и температура водяного конденсата (

), выше которой нельзя нагреть охлаждающий кислород (сокращается располагаемый хладоресурс криогенного окислителя);
- повышение параметра T_от благоприятно сказывается на тепловом балансе системы охлаждения вследствие снижения

, однако получаемый выигрыш снижается из-за того, что температура отработавшего пара (T_2T) отдаляется от кривой начала конденсации;
- для снижения затрат энергии на привод водяного насоса и особенно для улучшения теплового баланса системы охлаждения необходимо снижать

, однако со снижением этого параметра падают кпд водяного насоса и кпд турбины;
- с увеличением величины π_т при неизменном параметре р_2T величина L_ад возрастает лишь до некоторого предела, после чего уменьшается; кроме того, возрастают затраты мощности на водяной насос и снижается кпд турбины (вследствие повышения начальной плотности пара и возможного появления в турбинном тракте конденсированной фазы).The complex nature of the graphs in figure 2 is due to the action of a number of conflicting factors. In particular:
- with a decrease in p _2T while maintaining the initial pressure (that is, with an increase in the triggered pressure drop across the turbine π _t ), the work L _hell increases with a concomitant decrease

, and the temperature of the exhaust steam (T _2T ) is reduced, which is beneficial for the overall mass characteristics of the heat exchanger-condenser; but the temperature of the water condensate also decreases (

) above which cooling oxygen cannot be heated (the available cold resource of the cryogenic oxidizer is reduced);
- an increase in the parameter T _from favorably affects the thermal balance of the cooling system due to a decrease

however, the gain is reduced due to the fact that the temperature of the exhaust steam (T _2T ) is moving away from the curve of the beginning of condensation;
- to reduce energy costs for driving a water pump and especially to improve the heat balance of the cooling system, it is necessary to reduce

however, with a decrease in this parameter, the efficiency of the water pump and the efficiency of the turbine fall;
- with an increase in π _t with a constant parameter p _{2T, the} value of L _hell increases only to a certain limit, and then decreases; in addition, the power consumption for the water pump increases and the efficiency of the turbine decreases (due to an increase in the initial vapor density and the possible appearance of a condensed phase in the turbine path).

The graphs in FIG. 2 determine the range of p _2T = 0.6-4.0 MPa as the most suitable for the operation of the liquid-propellant rocket engine with the Rankine cycle on water vapor:
- the specified range of p _2T corresponds to positive values of the parameter t _article , which allows not to fear freezing of water condensate when cooling the exhaust turbine steam;
- in the specified range p _2T are the minimum values of the parameter s _rel , that is, the best overall (and therefore mass) performance of the heat exchanger-condenser are achieved.

The optimal range p _2T = 0.6 - 4.0 MPa corresponds to the operation of the water pump in the region of a small or even negative additional (over saturated vapor pressure) cavitation support at the pump inlet. In this case, the input values of the absolute pressure of the working fluid and temperature can significantly exceed the usual values for pumps.

, что грозит замерзанием конденсата, и во-вторых, снижается располагаемый хладоресурс жидкого окислителя, что приводит к быстрому возрастанию габаритов и массы теплообменника-конденсатора, а вскоре и к невозможности его реализации. Увеличение p_2T относительно указанного диапазона (р_2T>4,0 МПа) также нежелательно ввиду возрастания

, что приводит к чрезмерному увеличению габаритов и массы теплообменника-конденсатора и других функциональных элементов парожидкостного контура.A decrease in p _2T relative to the indicated range (p _2T <0.6 MPa) is undesirable for two reasons: first, the temperature drops

, which threatens to freeze the condensate, and secondly, the available refrigerant liquid oxidizer decreases, which leads to a rapid increase in the dimensions and mass of the heat exchanger-condenser, and soon to the impossibility of its implementation. An increase in p _2T relative to the indicated range (p _2T > 4.0 MPa) is also undesirable due to the increase

, which leads to an excessive increase in the dimensions and mass of the heat exchanger-condenser and other functional elements of the vapor-liquid circuit.

When choosing a specific value of p _2T within the specified range, one should take into account, in particular, the fact that at lower values of π _{t the} high efficiency of the turbine is achieved with a smaller number of steps, which simplifies and facilitates the design of the turbine and the entire TNA.

As illustrated by the example below, adhering to the specified range of p _2T also ensures a high level of p _k .

An example embodiment of the invention: LRE fuel oxygen-kerosene with a thrust of 1.5 MN:
oxidizer consumption through the engine - 350 kg / s;
fuel consumption through the engine - 135 kg / s;
closed-circuit water flow rate - 55 kg / s;
steam temperature at turbine inlet / outlet - 1000 K / 650 K;
steam pressure at the turbine inlet / outlet - 31 / 2.45 MPa;
water enters the pump with a pressure of 2.0 MPa at a temperature of 210 ^o C;
the working surface of the heat exchanger-condenser is 55 m ² with a design weight of 250 kg;
minimum wall temperature of the heat exchanger-condenser from the side of the water working fluid 50 ^o C;
p _to = 30 MPa.

Obtained in a specific example, the value of t _ct = 50 ^o With eliminates the freezing of water condensate, and the realized value of p _to = 30 MPa three times exceeds this parameter for the prototype method. So, the expected technical result is confirmed.

Claims (1)

 The method of operation of a liquid rocket engine with a closed vapor-liquid circuit in a turbopump supply system in which the working steam for a turbine is obtained from water using heat from the products of rocket fuel combustion, and the generated steam is cooled using an oxidizer coolant, and the condensate is returned to the water pump , characterized in that the steam is triggered to a pressure of 0.6-4.0 MPa.

Images