JP5909809B2 - Thrust control device and thrust control method for liquid rocket engine - Google Patents

Description

  The present invention relates to a thrust control device and a thrust control method for stabilizing a thrust when the thrust of a liquid rocket engine is reduced.

という関係がある。ここで、Ｔはロケットエンジンの推力、Ｃ_Fは推力係数、ｐ_cは燃焼圧、Ａ_tはノズルスロート面積、ｃ^*はロケットエンジンの特性速度、

は推進薬（燃料、酸化剤）の噴射流量である。 In liquid rocket engines

There is a relationship. Here, T is the rocket engine thrust, C _F is thrust coefficient, p _c is the combustion pressure, A _t is the nozzle throat area, c ^* is characteristic velocity of the rocket engine,

Is the injection flow rate of propellant (fuel, oxidant).

  As is clear from the above relationship, the thrust, combustion pressure, and propellant flow rate are almost proportional.When reducing the thrust, thrust is reduced by reducing the propellant flow rate and combustion chamber combustion pressure from the rated operating conditions. Reduction control is performed.

で支配される。ここで、Ｃは噴射オリフィスの流量係数、Ａは噴射オリフィスの断面積、ρは噴射流体の密度、ｐ_iは噴射オリフィス背部の噴射圧、ｐｃは燃焼室の燃焼圧である。これから、以下のように燃焼圧と噴射差圧（噴射圧と燃焼圧との差、Δｐ_ic）との関係を導くことができ、

となり、燃焼圧に対する噴射差圧の比は燃焼圧に比例することがわかる。これは、例えば推力を１０％に絞ると、噴射差圧と燃焼圧との比も定格時の１０％になることを意味する。非特許文献１（Crocco）によれば、低周波振動燃焼（いわゆる「チャギング」）の発生に関して、噴射差圧と燃焼圧との比

は重要なパラメータであり、この値が大きければ安定化し、この値が小さくなれば不安定化するとされている。従って、推力を深く絞った低減制御を行うと低周波振動燃焼（チャギング）のリスクが増加することが示唆される。 The flow rate of the propellant flowing into the combustion chamber through the injection orifice of the rocket engine is the following formula derived from Bernoulli's formula

Dominated by. Here, C is the flow coefficient of the injection orifice, A is the cross-sectional area of the injection orifice, ρ is the density of the injection fluid, p _i is the injection pressure at the back of the injection orifice, and pc is the combustion pressure in the combustion chamber. From this, the relationship between combustion pressure and injection differential pressure (difference between injection pressure and combustion pressure, Δp _ic ) can be derived as follows:

Thus, it can be seen that the ratio of the injection differential pressure to the combustion pressure is proportional to the combustion pressure. This means that, for example, when the thrust is reduced to 10%, the ratio between the injection differential pressure and the combustion pressure is also 10% of the rated value. According to Non-Patent Document 1 (Crocco), the ratio of the injection differential pressure to the combustion pressure with respect to the occurrence of low frequency vibration combustion (so-called “chugging”).

Is an important parameter. It is said that if this value is large, it will become stable, and if this value becomes small, it will become unstable. Therefore, it is suggested that the risk of low-frequency vibration combustion (chugging) increases if reduction control with deep thrust is performed.

とおき、噴射差圧と流量の関係式の微分を取ると、燃焼圧力の変動割合に対する流量変動割合は

となり、燃焼圧変動割合の半分に比例し、燃焼圧の変化に対しては、燃焼圧を小さくした時に逆比例して大きくなる。 Dividing flow rate and pressure into steady and variable terms

Then, taking the derivative of the relational expression between injection differential pressure and flow rate,

Thus, it is proportional to half of the fluctuation rate of the combustion pressure, and the change in the combustion pressure increases in inverse proportion when the combustion pressure is reduced.

  When reducing the thrust of the rocket engine, the combustion pressure and the propellant flow rate become smaller in proportion to the thrust. Therefore, the above equation shows that when the combustion pressure is lowered to reduce the thrust, the influence on the flow rate fluctuation ratio is amplified with respect to the combustion pressure fluctuation ratio of the same order. This means that if the thrust is controlled deeply in the reduction region, an unstable result in which the flow rate fluctuates very sensitively with respect to the fluctuation rate of the combustion pressure is produced.

  In Non-Patent Document 1 (Crocco), the cause of low-frequency vibration combustion (chugging) is caused by the combination of such flow rate fluctuations and time delays in the combustion process of the combustion chamber. In this case, the fact that the flow rate fluctuation is amplified even with a slight fluctuation in the combustion pressure indicates that the risk of low-frequency vibration combustion (chugging) increases from this aspect as well.

  Low-frequency vibration combustion (chugging) itself does not lead to combustion chamber destruction such as combustion chamber burning due to high-frequency vibration combustion, but is not acceptable from the viewpoint of limiting acceleration to the payload. It is necessary to take a solution for this problem in the reduction control.

Conventional techniques for dealing with the above problems can be summarized as follows.
(A) “Use of cavitating venturi inserted in propellant piping”
If the downstream pressure of the Venturi tube continues to decrease, the flow rate continues to increase and the pressure drops at the venturi throat. When the pressure in the throat reaches the saturated vapor pressure at the temperature of the liquid, no matter how much the downstream pressure continues to be reduced, the throat generates vapor without the pressure falling below the saturated vapor pressure, and the flow rate is the upstream pressure of the venturi nozzle. Maintain a constant flow according to the differential pressure determined by the saturated vapor pressure of the throat and throat. This is called cavitating venturi.

  Non-Patent Document 2 (Randall) describes the characteristics of a cavitating venturi in the case of water, and states that if this is applied to a rocket, a constant flow rate can be maintained without complicated control. In order to change the flow rate, a method of changing the passage area by inserting a pintle into the throat is introduced. The cavitating venturi is supposed to function in a state where the total diffuser pressure is 85% or less of the total upstream pressure and higher than the vapor pressure of the liquid. For liquid rocket engine applications, a cavitating venturi is inserted into the propellant supply pipe.

  In Patent Document 1, it is assumed that a cavitation orifice is placed in a propellant supply line for a two-component rocket engine and is used for stabilization of combustion at the time of reduced thrust of a tank type engine. The cavitating venturi of Non-Patent Document 2 (Randall) proposes that a pintle mechanism is inserted into the throat of the venturi to cope with a wide flow range. In Patent Document 1, a large flow rate and a small flow rate are branched so that precise adjustment of the pintle mechanism is not required, and a cavitation orifice is inserted into the small flow rate side pipe, and the upstream side valve is used as a pipe. The main effect of the patent is explained in terms of simplifying the response to changes in flow rate by switching the path. In the embodiment, a tank type engine of hydrazine / NTO combination is disclosed.

  By the way, when the cavitating venturi is inserted in the middle of the pipe line upstream of the combustion chamber of the rocket engine, the flow rate in the upstream portion of the cavitating venturi is determined by the upstream pressure and the throat area, and is stable in time. On the other hand, in the case of thrust reduction control, the injection differential pressure parameter in the injector remains very small, and the recovery pressure in the diffuser section downstream of the cavitating venturi shows a value almost close to the combustion pressure. Interlocks with pressure. This means that the effect of suppressing the chugging due to the selection of an appropriate injection differential pressure parameter as described in Non-Patent Document 1 (Crocco) is small. Since there are bubbles due to cavitation downstream of the throat of the cavitating venturi, compressive behavior and fluid accumulation effects are assumed. In particular, it is known that a cryogenic fluid has a more compressive behavior than a general incompressible liquid stored at room temperature. Therefore, the downstream piping portion of the cavitating venturi allows the flow rate and the pressure to be changed due to the influence of the fluctuation of the combustion pressure, and the effect is considered to be limited from the viewpoint of suppressing the chagging. From the viewpoint of complete prevention of chagging at the time of reduced thrust based on the theory of Non-Patent Document 1 (Crocco), the combustion pressure fluctuation is changed so that the effective injection differential pressure does not become small in the injection orifice of the injector. It is necessary to take a method that does not affect

(B) “Rating design for selecting excessive injection differential pressure”
As is clear from the above formula (1) describing the flow rate of the propellant passing through the injection orifice, the required injection differential pressure is proportional to the square of the flow rate. The injection differential pressure decreases in proportion to the square of the ratio. Since (injection differential pressure) / (combustion pressure), which is known as a parameter related to the occurrence of chagging, is proportional to the combustion pressure, this parameter also decreases as the combustion pressure decreases, and it becomes easier to induce chagging.

  In order to avoid the occurrence of chagging, the injection differential pressure is set to 10 to 15% of the combustion pressure in the normal rating point design. For engines with thrust control, the simplest method to avoid the risk of chagging during reduced thrust is set so that (injection differential pressure) / (combustion pressure) does not become too low at low combustion pressure during reduced thrust. However, the rated point is designed so that it is intentionally excessive with respect to the normal injection differential pressure. For example, if the injection differential pressure parameter when the thrust is reduced to 10% is 10%, the injection parameter at the rated thrust is 100%, that is, an injection differential pressure equal to the combustion pressure is required. Since this does not hold as an actual engine system, the injection differential pressure is selected within a compromise range.

  In the CECE example of Non-Patent Document 3 (Giuliano), as shown in FIG. 2 of the same document, the injector is designed by such a method. Such measures reduce the risk of chagging during reduced thrust, but the required pump discharge pressure at the time of rating is higher than when no reduced thrust control is performed, especially in closed-cycle engine systems such as CECE. In view of the power balance of the turbo pump, a pump discharge pressure higher than a simple increase in injection differential pressure is required.

JP-A-6-280680

L. Crocco and Sin-I Cheng, Theory of Combustion Instability in Liquid Propellant Rocket Motors ,. Butterworths Scientific Publications, 1956, p44. L.N.Randall, Rocket Applications of the Cavitating Venturi, ARS J., Jan-Feb, 1952, pp28-31. Victor J. Giuliano, et.al., CECE: Expanding the Envelope of Deep Throttling Technology in Liquid Oxygen / Liquid Hydrogen Rocket Engines for NASA Exploration Missions, AIAA 2010-6724, 2010.

  As described above, conventional techniques include the insertion of a cavitating venturi in the propellant supply pipe upstream of the combustion chamber and the setting of a higher injection differential pressure parameter at the rated design point. It is not a reliable solution.

  The present invention has been made in view of the above situation, and in the thrust control of a rocket engine, when operating in the range of reduced thrust, it is due to an extreme decrease in the injection differential pressure parameter that is normally generated. An object of the present invention is to provide a thrust control device and a thrust control method for a liquid rocket engine, which can greatly reduce the risk of occurrence of chagging.

  In a thrust control device for a liquid rocket engine that injects a liquid propellant into a combustion chamber through an injection orifice, the first invention includes a temperature control means that controls the temperature of the liquid propellant on the upstream side of the injection orifice, During engine thrust reduction control, the temperature of the liquid propellant is raised by the temperature control means so that the saturated vapor pressure of the liquid propellant is higher than the combustion pressure required for the reduction control.

  In the first invention, the temperature control means is a heat exchanger that raises the temperature of the liquid propellant by using heat exchange with the combustion gas in a cooling jacket provided around the combustion chamber as a heat source. Can do. Moreover, the said temperature control means shall supply heat to the said liquid propellant from the high enthalpy gas flow used for the turbine drive of the turbo pump provided in the engine. Furthermore, the temperature control means increases the pump work per unit flow rate of the liquid propellant by recirculating from the discharge side of the turbo pump provided in the engine to the suction side, thereby raising the temperature of the liquid propellant It can be.

  According to a second aspect of the invention, in the thrust control method for a liquid rocket engine that controls thrust reduction of a liquid rocket engine that injects the liquid propellant into the combustion chamber through the injection orifice, the temperature of the liquid propellant is controlled upstream of the injection orifice. Temperature control means is provided, and the temperature of the liquid propellant is raised by the temperature control means so that the saturated vapor pressure of the liquid propellant is higher than the combustion pressure required for the reduction control during thrust reduction control of the engine. It is characterized by that.

  In the second invention, the temperature control means is a heat exchanger that raises the temperature of the liquid propellant by using heat exchange with the combustion gas in a cooling jacket provided around the combustion chamber as a heat source. Can do. Moreover, the said temperature control means shall supply heat to the said liquid propellant from the high enthalpy gas flow used for the turbine drive of the turbo pump provided in the engine. Furthermore, the temperature control means increases the pump work per unit flow rate of the liquid propellant by recirculating from the discharge side of the turbo pump provided in the engine to the suction side, thereby raising the temperature of the liquid propellant It can be.

(A) It is the schematic of the liquid rocket engine in embodiment of this invention, (b) expands and shows the injection orifice part enclosed with the dashed-dotted line of (a). It is the graph which showed roughly the relationship between the saturation temperature and saturation pressure of a liquid propellant. It is the figure which showed the influence on the injection flow fluctuation (vertical axis) by the combustion pressure fluctuation (horizontal axis) at the time of using the injection orifice by the conventional design. It is the figure which showed the saturated vapor pressure characteristic of liquid oxygen. It is the graph which showed the mode of the pressure before and behind the injection orifice at the time of liquid oxygen temperature being 120K and carrying out reduction control of thrust. It is the graph which showed the mode of the pressure before and behind the injection orifice at the time of liquid oxygen temperature setting to 130K and carrying out low region control of thrust. (A) is the schematic which showed the Example which applied this invention to the rocket engine of the gas generator cycle system which uses liquid oxygen / liquid hydrogen as a propellant, (b) is enclosed with the dashed-dotted line of (a) It is the figure which expanded and showed the part. It is the schematic which showed the example at the time of applying a prior art to the rocket engine of a gas generator cycle system which uses liquid oxygen / liquid hydrogen as a propellant.

  Hereinafter, embodiments of the present invention will be described. First, the points of the present embodiment will be briefly summarized.

  The method of the present embodiment inserts a cavitation venturi in the upstream pipe of the injector and indirectly suppresses the fluctuation of the liquid propellant flow rate, but generates cavitation inside the injection orifice of the injector, The point is that even if a pressure fluctuation occurs in the downstream combustion chamber, the influence of the injection differential pressure fluctuation does not affect the flow fluctuation. In other words, the flow of liquid propellant is made to choke at the injection orifice. Here, the so-called injection orifice has a venturi shape when pressure is recovered, but it is not always necessary to recover the pressure as a venturi shape, and the injection orifice 6 has a nozzle shape as shown in FIG. In this way, the nozzle-shaped orifice functions sufficiently, and in order to process a large number of injection orifices, it is easier to process and the manufacturing cost can be reduced if the nozzle shape is used instead of the venturi shape.

  As a specific method, the pressure upstream of the injection orifice is set so that cavitation occurs at a predetermined flow rate and the flow velocity of the injection orifice throat. An important point in this case is to set the saturated vapor pressure at which the cavitation of the propellant is generated to be higher than the combustion pressure necessary for the thrust for performing the reduction control. The suppression policy for low-frequency vibration combustion based on the theory of Non-Patent Document 1 (Crocco) is to increase the ratio of the injection differential pressure to the combustion pressure. It means that the ratio of becomes smaller. Physically, the limit is that there is no injection flow rate fluctuation with respect to combustion pressure fluctuation. By utilizing the occurrence of cavitation in the injection orifice, the fact that there is no fluctuation in the injection flow rate in the injection orifice, regardless of the fluctuation in the combustion pressure downstream of the injection orifice, is equivalent to this extreme state, so that the combustion is stabilized. .

  The setting of the saturated vapor pressure higher than the combustion pressure is the point of the present method. On the other hand, as shown in FIG. 2, a normal liquid propellant shows a single saturated vapor pressure value according to a saturated vapor pressure characteristic curve depending on each storage temperature, and therefore does not necessarily correspond to an arbitrary combustion pressure. The saturated vapor pressure is not suitable. Therefore, by controlling the temperature of the propellant with a heat exchanger provided separately so that an appropriate saturated vapor pressure can be obtained with respect to the combustion pressure, the injection orifice can be controlled against the desired thrust or combustion pressure. It becomes possible to obtain the choke characteristics of the liquid propellant.

In order to raise the temperature of the liquid propellant by the action of heat exchange, as a heat source and heating method incorporated in the engine system,
1. A heat exchanger using a high enthalpy gas flow (branching from the upstream of the turbine before or after driving the turbine) for use in driving a turbo pump turbine;
2. Heat exchange from the combustion gas in a cooling jacket around the combustion chamber,
3. A method of increasing the pump fluid temperature by increasing the pump work per unit of liquid propellant flow by recirculating from the discharge side of the turbo pump to the suction side,
and so on.
With either of these methods, any fluctuation in combustion pressure will not affect the flow rate of the liquid propellant through the injection orifice, so there will be a chagging that causes the combustion pressure fluctuation combined with the flow fluctuation. Disappear.

By doing the above,
1. When controlling thrust, the liquid propellant can have a choke characteristic at the injection orifice of the engine combustion chamber, and the fluctuation in flow rate relative to the fluctuation in combustion pressure can be completely suppressed. The risk of frequency vibration combustion can be eliminated.
2. By controlling the temperature of the liquid propellant to control the saturated vapor pressure that generates cavitation, the range of control regarding the operating point of the propellant cavitation is expanded with respect to the desired combustion pressure.
3. By reducing the injection differential pressure at the rated design point to reduce the risk of low-frequency vibration combustion in low-frequency thrust control, the design is designed to increase the discharge pressure of the turbo pump at the rated time. However, this system disadvantage can be avoided by using a method of generating cavitation controlled by the injection orifice, which causes performance loss or engine weight increase.

Next, this embodiment will be described more specifically. As a point,
1. In the injection orifice of the injector that injects the liquid propellant into the combustion chamber of the rocket engine, the flow of the liquid propellant at the throat part becomes high speed, and when the pressure drops and reaches the saturated vapor pressure, cavitation occurs, and more Therefore, even if the pressure in the downstream combustion chamber, which is set lower than the saturated vapor pressure, fluctuates using the phenomenon that the flow rate does not change, the liquid propellant that passes through the injection orifice does not change. Ensure that the flow rate does not fluctuate.
2. If a special method is not used when controlling the thrust of the rocket engine to a low range, the flow rate of the liquid propellant will fluctuate significantly when the differential pressure at the injection orifice decreases and the pressure in the combustion chamber changes. Thus, the occurrence of low-frequency combustion vibration is prevented by utilizing the fact that the flow rate fluctuation of the liquid propellant in the injection orifice does not occur by using the above method 1.
3. By controlling the temperature using the saturated vapor pressure characteristics of the liquid propellant, when controlling the thrust reduction, the liquid propellant is controlled to have an appropriate saturated vapor pressure with respect to the working combustion pressure, Applicable to a wide range of liquid propellant types and working combustion pressure.
4). As a heat source to raise the temperature of the liquid propellant, heat exchange with the gas that drives the turbo pump, heat exchange with the cooling jacket of the combustion chamber, enthalpy increase due to pump work due to reflux from the turbo pump discharge side to the suction side, gas generator , A pre-burner, heat exchange with gas from a separate combustor, or a combination thereof.

  Next, an example in a liquid hydrogen / liquid oxygen engine will be described. FIG. 3 shows the influence on the injection flow rate fluctuation (vertical axis) by the combustion pressure fluctuation (horizontal axis) when the injection orifice of the conventional design is used. As can be seen from this figure, the fluctuation of the injection flow rate with respect to the fluctuation of 1% of the combustion pressure is the same order as 2.4% when the thrust is 100%, but 24% when the thrust is 10%. Further, when the combustion pressure fluctuation reaches 4%, the flow fluctuation reaches 100%. Therefore, in the low-frequency control of thrust, the necessity of measures for suppressing propagation to flow rate fluctuation due to combustion pressure fluctuation is obvious.

  In a liquid hydrogen / liquid oxygen engine, hydrogen is usually injected in the state of a gas that is used for cooling the combustion chamber and has an increased enthalpy, and there is no particular problem even under low combustion pressure conditions. When the thrust is controlled to be reduced by this engine, fluctuations in the flow rate of liquid oxygen with respect to fluctuations in combustion pressure become a problem, so cavitation is generated at the injection orifice on the liquid oxygen side. Specifically, when the combustion pressure is lower than the rated conditions required for thrust control, the liquid oxygen is heated by heat exchange or the like so that the saturated vapor pressure of cavitation becomes higher than the combustion pressure. The injection pressure is adjusted so that a predetermined saturated vapor pressure is obtained at a predetermined flow rate at the contracted flow portion of the injection orifice. The saturated vapor pressure characteristics of liquid oxygen are as shown in FIG.

  The storage temperature of the liquid oxygen in a heat insulating container under 1 atm is 90K which is the boiling point under atmospheric pressure. In rockets, the propellant tank is pressurized to several atmospheres due to consideration of cavitation against pump suction, and the pump inlet is approximately 95K due to inflow of enthalpy by pressurized gas and heat inflow from piping. Flows in. In order to work with the pump, the temperature is increased by an amount corresponding to the work amount per unit mass. In the low-frequency control of the thrust that exerts the function of the present invention, since the combustion pressure is set low, the pump discharge pressure is also low, and the temperature rise due to pump work is small.

As an example, in an engine having a rated combustion pressure of 6 MPa, when the temperature of liquid oxygen is increased by 25 K by heat exchange to 120 K, the saturated vapor pressure is as shown in FIG. As shown in Table 1, it becomes 1.022 MPa, and when the temperature of liquid oxygen is 130 K, it becomes 1.749 MPa. FIG. 5 shows the state of pressure before and after the injection orifice when the thrust is controlled to be reduced in this case, and FIG. 6 shows the case where the temperature is raised to 130K. In these figures, the horizontal axis indicates the flow rate reduction rate, which is the same as the thrust reduction rate. 2: Pc is the combustion pressure, which is proportional to the flow rate and proportional to the thrust. 1: Pinj (Norm) indicates the injection pressure upstream of the injection orifice at each flow rate when the present invention is not used, and approaches the combustion pressure Pc as the thrust decreases.

  The saturation vapor pressure at the throat of the injection orifice when using the present invention is 4: Ps (Ts = 120K), 4: Ps (Ts = 130K). 3: Pinj (Ps, keep) uses the present invention to reach the saturated vapor pressure at the throat of the injection orifice and generate the cavitation, and the injection pressure upstream of the injection orifice necessary to flow the reduced flow rate. Is shown.

  When the temperature of liquid oxygen is 120 K, the saturated vapor pressure is 1.022 MPa. From FIG. 5, when the thrust is reduced to 10%, the injection differential pressure before and after the injection orifice in the cavitation generation state is (Pinj (Ps, keep) −Pc) = 0.444 MPa. Since Pc is about 0.6 MPa, that is, (Pinj (Ps, keep) −Pc) /Pc=0.74 can be secured. Even when the thrust is reduced to 15%, (Pinj (Ps, keep) −Pc) /Pc=0.19 and an injection differential pressure corresponding to the rated point can be secured. Actually, when the thrust is reduced to 15%, even if the combustion pressure fluctuates upward by 10%, the flow rate does not fluctuate because the downstream side of the contracted portion of the injection orifice is below the saturated vapor pressure. As long as the flow rate does not fluctuate, even if a combustion time delay occurs, there is no risk of occurrence of chagging because a phenomenon in which vibration is amplified due to a phase shift does not occur.

  As shown in FIG. 6, when the temperature of the liquid oxygen is further increased to 130 K, the saturated vapor pressure becomes 1.749 MPa. In this case, the flow rate can be stabilized by the present invention until the thrust is controlled to about 25%, and (Pinj (Ps, keep) −Pc) /Pc=0.26 can be secured. Even if the combustion pressure fluctuates upward by 10%, the flow rate does not vary at all because the downstream of the constricted flow portion of the injection orifice is below the saturated vapor pressure.

Next, examples of the self-combustible propellant engine will be described. As a specific example of the self-combustible propellant, a case where the storage temperature of hydrazine / NTO is an ambient temperature (25 ° C.) is taken up. There are several types of hydrazine, but their saturated vapor pressure at ambient temperature is 0.0019 MPa for N ₂ H ₄ , 0.0074 MPa for MMH, and 0.023 MPa for UDMH. On the other hand, the N ₂ O _4, is 0.12 MPa. The normal rated combustion pressure of a tank type engine using these propellants is 0.5 to 1 MPa. The combustion pressure during thrust control is 0.15 to 0.3 MPa for a 30% reduction thrust and 0.05 to 0.1 MPa for a 10% reduction thrust. At the time of reduced thrust, the propellant in which the injector orifice functions as a cavitation orifice is only N ₂ O ₄ at 10% reduced thrust.

Now, when the present invention is applied, the saturation vapor pressure becomes 0.062 MPa when the temperature is raised to 50 ° C. in the case of UDMH, and the present invention can be applied if the combustion pressure is up to about 0.05 MPa. In the case of MMH, when the temperature is raised to 79 ° C., the saturated vapor pressure becomes 0.071 MPa, and the present invention can be applied to a combustion pressure of about 0.05 MPa. N ₂ O ₄ has a saturated vapor pressure of 0.12 MPa at room temperature (25 ° C.), so it can be used as it is up to a combustion pressure of 0.1 MPa. However, when the temperature is raised to 30 ° C., the saturated vapor pressure becomes 0.15 MPa, The present invention can be applied if the combustion pressure is up to about 0.12 MPa.

  The above calculation corresponds to the case where the injection orifice is a nozzle. However, if the injection orifice is a venturi, the combustion pressure to be applied is further increased by recovering the pressure with the venturi, and the low frequency vibration combustion is performed. It is possible to expand the stable range for.

  In the above, the application to the rocket engine has been described. In addition, for example, in a boiler that burns liquid fuel such as heavy oil, a combustion device that sprays fuel through an injection valve can be used to change the load in a wide range. In a required apparatus, there is a risk that low-frequency vibration combustion is generated in the same manner as combustion in a rocket engine when load is reduced. In such a case, the risk can be eliminated by using the present invention.

  FIG. 7 shows an embodiment in which the present invention is applied to a gas generator cycle rocket engine using liquid oxygen / liquid hydrogen as a propellant. Liquid hydrogen flows into the LH2 pump from the engine interface and is pressurized, and then cools the walls of the thrust chamber. After that, most of the hydrogen is combusted with liquid oxygen in the thrust chamber and contributes to the generation of thrust. It is supplied to the gas generator and burned with a small amount of liquid oxygen to generate a gas that drives the turbine. This gas is discarded in the rated operation after driving the turbo pump of liquid hydrogen and liquid oxygen, but the thrust of When performing low-frequency control, it is dumped via a heat exchanger. When the amount of heat required by the heat exchanger is insufficient, the gas from the gas generator is mixed with the gas after driving the turbine, bypassing the turbine. Liquid oxygen flows into the LOX pump from the engine interface and is boosted. Then, in rated operation, it is supplied to the thrust chamber and the gas generator at a predetermined rate and used for combustion, but when performing thrust reduction control, The temperature is raised to a predetermined temperature by the heat exchanger and then supplied to the thrust chamber and the gas generator for combustion.

  The orifice on the liquid oxygen side of the injection element of the injector that injects liquid oxygen into the combustion chamber of the thrust chamber is in the form of a nozzle or a venturi, and liquid oxygen is injected without any phase change from the rated value in the mid-range thrust control. However, in thrust reduction control, as the injection pressure is lowered to a predetermined pressure, cavitation occurs, the pressure in the throat section maintains the saturated vapor pressure, and fluctuations in flow rate due to fluctuations in combustion pressure are eliminated. Although the present description of the embodiment is for a gas generator cycle, there is essentially no difference between engine cycles if there is a heat source to raise the liquid oxygen to a predetermined temperature.

  FIG. 8 shows an example in which the prior art is applied to a gas generator cycle rocket engine using liquid oxygen / liquid hydrogen as a propellant. The difference from the embodiment of the present invention is that a cavitating venturi is inserted at the position of the heat exchanger. Cavitation occurs in this cavitating venturi and aims to contribute to stabilization of the flow rate. However, although the flow rate upstream of the cavitating venturi can be stabilized, the effect of suppressing flow rate fluctuation is limited because the compressive behavior due to cavitation bubbles appears downstream. Therefore, the risk of low frequency vibration combustion cannot be removed.

  If the cavitating venturi is not used, the injection differential pressure during rated operation is simply designed to be high. For this reason, the difference in the engine system causes a loss of performance at the most frequently used rated operation and an increase in the weight of the turbo pump.

Claims (8)

In a thrust control device for a liquid rocket engine that injects liquid propellant into a combustion chamber through an injection orifice,
A temperature control means for controlling the temperature of the liquid propellant on the upstream side of the injection orifice;
When thrust reduction control of the engine, the saturated vapor pressure of the liquid propellant passes through the injection orifice Ri of higher than combustion pressure necessary for the reduction control, and so that cavitation occurs in the injection orifice, A thrust control apparatus for a liquid rocket engine, wherein the temperature of the liquid propellant is raised by the temperature control means.   2. The liquid according to claim 1, wherein the temperature control means is a heat exchanger that raises the temperature of the liquid propellant by using heat exchange with combustion gas in a cooling jacket provided around the combustion chamber as a heat source. A rocket engine thrust control device.   The thrust control apparatus for a liquid rocket engine according to claim 1, wherein the temperature control means supplies heat to the liquid propellant from a high enthalpy gas flow used for driving a turbine of a turbo pump provided in the engine.   The temperature control means raises the temperature of the liquid propellant by increasing the pump work per unit flow rate of the liquid propellant by returning from the discharge side of the turbo pump provided in the engine to the suction side. The thrust control apparatus for a liquid rocket engine according to claim 1. In a liquid rocket engine control method for controlling thrust reduction of a liquid rocket engine that injects a liquid propellant into a combustion chamber through an injection orifice,
A temperature control means for controlling the temperature of the liquid propellant is provided on the upstream side of the injection orifice, and the saturated vapor pressure of the liquid propellant when passing through the injection orifice is used for the reduction control during thrust reduction control of the engine. Ri a higher than required combustion pressure, and so that cavitation occurs in the spray orifice, thrust control method of the liquid rocket engine, characterized by heating the liquid propellant by the temperature control means.   The liquid according to claim 5, wherein the temperature control means is a heat exchanger that raises the temperature of the liquid propellant by using heat exchange with a combustion gas in a cooling jacket provided around the combustion chamber as a heat source. A rocket engine thrust control method.   6. The liquid rocket engine control method according to claim 5, wherein the temperature control means supplies heat to the liquid propellant from a high enthalpy gas flow used for driving a turbine of a turbo pump provided in the engine.   The temperature control means raises the temperature of the liquid propellant by increasing the pump work per unit flow rate of the liquid propellant by returning from the discharge side of the turbo pump provided in the engine to the suction side. A thrust control method for a liquid rocket engine according to claim 5.

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