JP7250304B2 - Propellant for gas generator for turbine drive and method for producing the same - Google Patents

Description

TECHNICAL FIELD The present invention relates to a propellant for a gas generator for driving a turbine and a method for producing the same.

A gas generator cycle air turboramjet engine (hereinafter referred to as a "GG-ATR engine") has a thrust-to-weight ratio higher than that of a conventional jet engine. there is For example, Non-Patent Document 1 discloses a GG-ATR engine in which ethanol as a fuel and liquid oxygen as an oxidant are separately supplied to a gas generator to generate combustion gas.

Ryojiro Minato, System operating characteristics of GG-ATR engine, Muroran Institute of Technology Aerospace System Research Center Annual Report, Vol. 2014, p. 20-22, 2015

The GG-ATR engine's propellant, which consists of ethanol and liquid oxygen, is non-toxic and can provide high thrust. However, since liquid oxygen is a cryogenic fluid with a boiling point of about -180°C, it takes a lot of work to fill the tank of the GG-ATR engine with liquid oxygen. There is also a need for a propellant that is safer and storable at ambient temperature to facilitate operation of the GG-ATR engine. Such problems are not limited to propellants for GG-ATR engines, but also exist for propellants for other turbine-driving gas generators.

SUMMARY OF THE INVENTION It is an object of the present invention to provide a propellant for a gas generator for driving a turbine, which is easy for users to handle and can be stored even at room temperature, and a method for producing the same. .

In order to achieve the above object, a propellant for a turbine-driven gas generator according to a first aspect of the present invention comprises:
A mixture of alcohol and nitromethane is mixed so that the weight of the nitromethane is greater than the weight of the alcohol .

The alcohol may be at least one selected from the group consisting of ethanol, methanol, propanol, ethylene glycol, propylene glycol and glycerin.

In order to achieve the above object, a propellant for a turbine-driven gas generator according to a second aspect of the present invention comprises:
A mixture of alcohol and nitromethane, wherein the alcohol is ethanol, and the weight ratio of the nitromethane to the ethanol is 0.5 . 8 to 2 . 5.

Methylamine may be added.

In order to achieve the above object, a method for producing a propellant for a turbine-driven gas generator according to a third aspect of the present invention comprises:
A method for producing a propellant for the turbine driving gas generator, comprising:
a weighing step of respectively weighing the alcohol and the nitromethane ;
a mixing step of mixing the alcohol and the nitromethane weighed in the weighing step;
including.

Advantageous Effects of Invention According to the present invention, it is possible to provide a propellant for a gas generator for driving a turbine, which is easy for users to handle and can be stored even at room temperature, and a method for producing the propellant.

1 is a diagram showing a configuration of a gas generator cycle air turboramjet engine according to an embodiment of the present invention; FIG. FIG. 3 is a diagram showing physical property values of ethanol and nitromethane; 1 is a flow chart showing the flow of a propellant manufacturing method according to an embodiment of the present invention; 4 is a diagram showing analysis conditions for combustion simulation in Example 1. FIG. 4 is a graph showing the relationship between specific thrust and gas generator combustion temperature in Example 1. FIG. 4 is a graph showing the relationship between Isp and gas generator combustion temperature in Example 1. FIG. 4 is a graph showing the relationship between ram combustor temperature and gas generator combustion temperature in Example 1. FIG. 4 is a graph showing the relationship between the weight ratio of nitromethane to ethanol and the combustion temperature of the gas generator in Example 1. FIG. 9 is a graph showing the relationship between specific thrust and gas generator combustion temperature in Example 2. FIG. 7 is a graph showing the relationship between Isp and gas generator combustion temperature in Example 2. FIG. 7 is a graph showing the relationship between ram combustor temperature and gas generator combustion temperature in Example 2. FIG.

DETAILED DESCRIPTION OF THE INVENTION A propellant for a turbine-driving gas generator and a method for producing the propellant according to an embodiment of the present invention will now be described in detail with reference to the drawings. In each drawing, the same reference numerals are given to the same or equivalent parts.

In the embodiment, a case where the propellant for the turbine driving gas generator and the manufacturing method thereof is applied to a gas generator cycle air turbo ramjet (GG-ATR) engine will be described as an example. However, the present invention is not limited to this case.

The configuration of the GG-ATR engine will be described with reference to FIG. The GG-ATR engine is a type of ramjet engine. A ramjet engine is a jet engine that obtains thrust by blowing and burning fuel against compressed air compressed by ram pressure generated in an air intake.

A ramjet engine is suitable for supersonic flight of about Mach 3 to 5, and since it does not require air compression by a compressor, it has a simpler structure than a turbojet engine. A ramjet engine, on the other hand, requires a separate propulsion system in order to allow the airframe to reach a predetermined speed range where compression of air by the compressor is unnecessary. An air turbo ramjet engine is equipped with a mechanism equivalent to a turbojet engine, such as a compressor, inside the ramjet engine so that it functions as a turbojet engine until the airframe reaches a predetermined speed range. It is configured.

The GG-ATR engine is configured to supply combustion gas generated by burning a propellant in a gas generator to a turbine, mix the combustion gas and compressed air in the turbine, and further burn it.

FIG. 1 is a longitudinal sectional view of a part of a GG-ATR engine according to an embodiment. The GG-ATR engine 1 comprises a tank 11, an air intake 12, a compressor 13, a gas generator 14, a turbine 15, a ram combustor 16 and a nozzle 17.

The tank 11 is a tank that is connected to the gas generator 14 via a pipe or the like and stores the propellant to be supplied to the gas generator 14 . The tank 11 may be, for example, a single tank if the propellant is a one-liquid system, but if it is a two-liquid system propellant that requires a fuel and an oxidizer, a fuel tank and an oxidizer tank may be used. must be kept separately.

The air intake 12 is arranged on the tip side of the GG-ATR engine, and supplies air taken in from the front to the compressor 13 .

The compressor 13 is arranged behind the air intake 12 , and a turbine 15 mechanically connected to the compressor 13 is driven to compress air taken in from the air intake 12 and supply the air to the turbine 15 . The compressor 13 includes a rotating shaft connected to the turbine 15 and a plurality of blades radially extending from the outer peripheral surface of the rotating shaft. Air taken in from the air intake 12 is compressed by the rotating blades and discharged rearward.

The gas generator 14 is connected to the tank 11 so as to be supplied with propellant from the tank 11 , burns the propellant supplied from the tank 11 , and supplies the generated high-temperature, high-pressure combustion gas to the turbine 15 .

The turbine 15 is arranged behind the compressor 13 and driven by the high-temperature, high-pressure combustion gas supplied from the gas generator 14 to drive the compressor 13 . The turbine 15 includes a rotating shaft connected to the compressor 13 and a plurality of heat-resistant blades radially extending from the outer peripheral surface of the rotating shaft. The rotating shaft of the turbine 15 is connected so that the rotating shaft of the compressor 13 and the central axis are aligned. The heat-resistant temperature of the heat-resistant blades forming the turbine 15 is about 1100K. The combustion gas supplied from the gas generator 14 is injected to the heat-resistant blades of the turbine 15 to drive the turbine 15 , and then discharged to the rear of the turbine 15 .

The ram combustor 16 is arranged behind the turbine 15 and mixes the combustion gas discharged from the turbine 15 with the compressed air supplied from the compressor 13 to further burn the high-temperature, high-pressure combustion gas.

The nozzle 17 is arranged behind the ram combustor 16 and injects combustion gas further combusted in the ram combustor 16 rearward. The GG-ATR engine generates thrust by injecting combustion gas from the nozzle 17 . The above is the configuration of the GG-ATR engine.

Next, the composition of the GG-ATR engine propellant according to the embodiment will be described. The propellant for the GG-ATR engine is a mixture of alcohol, the fuel, and nitromethane, the oxidant.

FIG. 2 shows physical property values of ethanol and nitromethane as examples of alcohols. As can be understood from the melting points and boiling points shown in FIG. 2, both nitromethane and ethanol are liquid at room temperature. Nitromethane has the property of being well mixed with alcohol molecules, and both can be stored in a mixed state. Therefore, by mixing alcohol and nitromethane in advance, even a two-component propellant can be treated substantially like a one-component propellant.

The propellant alcohol is preferably a lower alcohol, more preferably a primary alcohol. The propellant alcohol includes, for example, ethanol, methanol, propanol, ethylene glycol, propylene glycol, glycerin, etc. Ethanol, methanol and propanol are preferred, and ethanol is more preferred. Ethanol may be, for example, bioethanol obtained by fermenting and distilling biomass such as sugarcane and corn, other fermented ethanol, synthetic ethanol, and the like.

The alcohol of the propellant is not limited to being composed of a single composition, for example, a mixture of at least two alcohols from the group consisting of ethanol, methanol, propanol, ethylene glycol, propylene glycol, glycerin, etc. can be anything. When a plurality of types of alcohol are mixed, it is preferable to mix them so that the composition of C, H, and O is substantially the same as that of ethanol, for example.

The propellant, nitromethane, acts as an oxidant to oxidize the alcohol fuel. Nitromethane is insensitive to external impact and its explosive power is comparable to that of black powder. Nitromethane can be handled relatively easily because the risk of explosion is further reduced by mixing it with alcohol.

Nitromethane is easy to handle due to its relatively low toxicity. The median lethal dose (50%: LD50) of nitromethane is about 940 mg/kg. The LD50 is the dose at which half of the animals dosed die. Since the LD50 of nitromethane is 300 mg/kg or more, it does not fall under the Poisonous and Deleterious Substances Control Law and does not require special management.

Considering the heat resistance of the GG-ATR engine 1, the weight ratio of ethanol to nitromethane is, for example, in the range of about 0.8 to about 2.5 for 1.0 ethanol, and about 1.0 for nitromethane. It is preferably in the range of 4 to about 2.0, more preferably in the range of about 1.7 to about 1.9.

In the GG-ATR engine, the higher the combustion temperature of the gas generator 14 (hereinafter abbreviated as "GG combustion temperature"), the higher the propulsion efficiency. Since the heat-resistant temperature of the heat-resistant blades constituting the turbine 15 is about 1100K, it is preferable to set the weight ratio of ethanol and nitromethane so that the GG combustion temperature is 1100K. In order to obtain a GG combustion temperature of about 1100K, the weight ratio of ethanol to nitromethane should be about 1.8 of nitromethane to 1.0 of ethanol.

The propellant for the GG-ATR engine may have a mixture of ethanol and nitromethane with a trace amount of methylamine added. Nitromethane alone has low ignitability, so adding a small amount of methylamine can improve the ignitability of the propellant. Preferably, the amount of methylamine is about 1/10 or less of ethanol by weight. More specifically, the weight ratio of ethanol to nitromethane to methylamine is preferably about 1.8 of nitromethane and about 0.1 of methylamine to 1.0 of ethanol.

Next, referring to FIG. 3, the flow of the manufacturing process of the propellant for the GG-ATR engine will be described. FIG. 3 is a flow chart showing the flow of the manufacturing process of the propellant for the GG-ATR engine. An example in which ethanol is selected as the alcohol of the propellant for the GG-ATR engine will be described below.

First, ethanol and nitromethane, which are materials constituting the propellant for the GG-ATR engine, are weighed (step S1). In step S1, ethanol and nitromethane are weighed so that the weight ratio of ethanol:nitromethane=1.0:1.8, for example.

Next, the ethanol and nitromethane weighed in step S1 are mixed in the container (step S2). More specifically, the process of step S2 may be performed by previously charging and mixing the ethanol and nitromethane weighed in step S1 into another container before charging into the tank 11. Alternatively, the process of step S2 may be executed by directly charging the ethanol and nitromethane weighed in step S1 into the tank 11 respectively.

When adding a small amount of methylamine to the propellant for the GG-ATR engine, the methylamine may be weighed in step S1 and mixed with ethanol and nitromethane in step S2. The above is the flow of the manufacturing process of the propellant for the GG-ATR engine.

EXAMPLES The present invention will be specifically described below with reference to Examples. However, the present invention is not limited to these examples.

(Example 1)
In this verification, an analysis model was created to analyze the combustion state of the GG-ATR engine, and a simulation of the combustion state when the target fuel was burned was performed using the analysis model. This combustion simulation utilizes chemical equilibrium calculation by the Gibbs free energy minimization method. In addition, research institutes such as the US National Aeronautics and Space Administration have also conducted simulations of the combustion of kerosene and ethanol in rocket engines through analysis using CEA (Chemical Equilibrium Application), etc., and the actual combustion results are well approximated. has been demonstrated.

In this verification, in addition to "ethanol-nitromethane", which is a mixture of ethanol C ₂ H ₅ OH and nitromethane CH ₃ NO ₂ of the present invention, hydrazine N ₂ H ₄ (monocomponent propellant) for comparison, “Hydroxy-terminated polybutadiene-ammonium perchlorate” (stationary propellant), which is a mixture of hydroxyl-terminated polybutadiene and ammonium perchlorate NH ₄ ClO ₄ , kerosene n-C ₁₂ H ₂₆ and hydrogen peroxide H ₂ O ₂ (concentration 60% or 100%) and "kerosene-hydrogen peroxide" (two-liquid system room temperature propellant) were respectively burned. Hereinafter, "hydroxyl-terminated polybutadiene-ammonium perchlorate" is abbreviated as HTPB-AP (Hydroxyl Terminated Poly Butadiene-Ammonium Perchlorate).

Hydrazine is self-igniting, but has an LD50 of 60 mg/kg and is extremely toxic. Since HTPB-AP is a solid propellant, it is easy to handle and does not require piping for fuel supply, but it is difficult to control thrust. Also, HTPB-AP is subject to regulations under the Explosives Control Law, so strict control is required. Kerosene-hydrogen peroxide is an explosion hazard at high concentrations of hydrogen peroxide.

When HTPB-AP is actually burned in a GG-ATR engine, it is necessary to mix metal powder such as aluminum and magnesium to stabilize combustion. On the other hand, in this verification, in order to simplify the calculation, the combustion simulation was performed without mixing the metal powder.

FIG. 4 shows engine parameter conditions in a combustion simulation. In the case of HTPB-AP, the combustion pressure of the gas generator 14 and the expansion ratio of the turbine 15 can be set higher than in the case of other propellants. The expansion ratio of was set to 16. Then, the combustion state was analyzed under the conditions shown in FIG.

The results of the combustion simulation are shown below with reference to FIGS. 5 to 8. FIG. FIG. 5 is a graph showing simulation results of specific thrust when each propellant is burned. The vertical axis is specific thrust (kgf sec/kg), and the horizontal axis is GG combustion temperature (K). Specific thrust is the thrust per airflow.

It can be seen that when the GG combustion temperature is 1100 K or less, the specific thrust of ethanol-nitromethane becomes higher than other propellants. Also, the ethanol-nitromethane specific thrust is maximized in the range of about 1000K to about 1050K. Considering the heat resistance of the heat-resistant blades of the turbine 15, the GG combustion temperature must be suppressed to about 1100K or less, so ethanol-nitromethane is the best in terms of thrust specificity in the practical range.

FIG. 6 is a graph showing simulation results of Isp when each propellant is burned. The vertical axis is Isp (sec) and the horizontal axis is GG combustion temperature (K). Isp, also called specific impulse, is the thrust generated per unit mass flow of fuel, and corresponds to the fuel consumption of an automobile engine.

It can be seen that when the GG combustion temperature is 1100 K or less, the Isp of ethanol-nitromethane is higher than that of HTPB-AP. Also, it can be seen that the Isp of ethanol-nitromethane is in good agreement with the Isp of N ₂ H ₄ when the GG combustion temperature is around 1100K. Considering the toxicity of N ₂ H ₄ , it can be judged that ethanol-nitromethane is highly superior.

FIG. 7 is a graph showing simulation results of ram combustor temperature when each propellant is burned. The vertical axis is the temperature (K) of the ram combustor 16 and the horizontal axis is the GG combustion temperature (K). As the ram combustor temperature increases, the propulsion efficiency of the GG-ATR engine improves.

In the case of ethanol-nitromethane, the ram combustor temperature reaches its maximum value when the GG combustion temperature is approximately 1100K, whereas in the case of HTPB-AP, the ram combustor temperature reaches its maximum value when the GG combustion temperature is 1300K. became. In the case of HTPB-AP, it is necessary to set the GG combustion temperature relatively high. It shows that the ram combustor temperature can be maximized.

FIG. 8 is a graph showing simulation results of GG combustion temperature when the weight ratio of ethanol and nitromethane is changed. The vertical axis is the ratio of nitromethane (O) to ethanol (F), and the horizontal axis is the GG combustion temperature (K).

Considering the heat resistance of the turbine 15, it is necessary to set the GG combustion temperature to about 1100K, and the O/F ratio at this time is about 1.8. Therefore, it can be understood that the weight ratio of ethanol to nitromethane is preferably ethanol:nitromethane=1.0:1.8.

(Example 2)
In this verification, using the same analysis model for analyzing the combustion state of the GG-ATR engine as in Example 1, a propellant composed of ethanol-nitromethane, methanol and nitromethane (hereinafter referred to as "methanol-nitromethane" ) and a propellant obtained by adding methylamine to ethanol-nitromethane (hereinafter referred to as "methylamine-added propellant").

The engine parameter conditions in each simulation are the same as for "ethanol-nitromethane" in FIG. The methylamine-added propellant contained fuel in a weight ratio of 90% ethanol and 10% methylamine.

FIG. 9, like FIG. 5, is a graph showing a simulation result of specific thrust when each propellant is burned. The vertical axis is specific thrust (kgf sec/kg), and the horizontal axis is GG combustion temperature (K). The specific thrust of the methylamine-added propellant closely matches that of ethanol-nitromethane. It can also be seen that the specific thrust of methanol-nitromethane is greater than that of ethanol-nitromethane in the range of GG combustion temperatures below about 1000K.

FIG. 10, like FIG. 6, is a graph showing the simulation results of Isp when each propellant is burned. The vertical axis is Isp (sec) and the horizontal axis is GG combustion temperature (K). The Isp of the methylamine-added propellant approximately matches that of ethanol-nitromethane. It can also be seen that the Isp of methanol-nitromethane is greater than that of ethanol-nitromethane in the range of GG combustion temperature below about 1050K.

FIG. 11, like FIG. 7, is a graph showing simulation results of the ram combustor temperature when each propellant is burned. The vertical axis is the temperature (K) of the ram combustor 16 and the horizontal axis is the GG combustion temperature (K). The ram combustor temperature for the methylamine-added propellant approximately matches that for ethanol-nitromethane. It can also be seen that the ram combustor temperature for methanol-nitromethane is greater than that for ethanol-nitromethane in the range of GG combustion temperatures below about 1000K.

From the above, it was confirmed that the same propulsion performance as ethanol-nitromethane can be achieved even if methylamine is added to ethanol-nitromethane to improve ignitability. It was also confirmed that even if the ethanol used as the propellant for the GG-ATR engine was replaced with methanol, propulsion performance almost equivalent to that of ethanol-nitromethane could be achieved.

Ethanol-nitromethane is also used as a combustion improver for small gasoline engines. However, the GG-ATR engine and the small gasoline engine differ greatly in the ethanol-nitromethane combustion mode. More specifically, in the GG-ATR engine, the methanol-nitromethane itself is burned in the absence of oxygen in the gas generator 14, whereas in the small gasoline engine, it is used to help burn the gasoline fuel. and combusted with gasoline in the presence of oxygen.

(Modification)
The present invention is not limited to the above embodiments, and modifications described below are possible.

In the above embodiment, the propellant containing alcohol and methylamine was used as the GG-ATR engine propellant, but the present invention is not limited to this. For example, propellants containing alcohol and methylamine may be used in other turbine driven gas generators. Turbine-driven gas generators may be used, for example, to drive turbopumps for rocket engines.

The above embodiments are examples, and the present invention is not limited to these, and various embodiments are possible without departing from the scope of the invention described in the claims. The components described in each embodiment and modifications can be freely combined. In addition, inventions equivalent to the inventions described in the claims are also included in the present invention.

1 GG-ATR engine 11 tank 12 air intake 13 compressor 14 gas generator 15 turbine 16 ram combustor 17 nozzle

Claims (5)

A propellant for a gas generator for driving a turbine, comprising a mixture of alcohol and nitromethane, wherein the weight of the nitromethane is greater than the weight of the alcohol. The alcohol is at least one selected from the group consisting of ethanol, methanol, propanol, ethylene glycol, propylene glycol and glycerin.
A propellant for a gas generator for driving a turbine according to claim 1. A mixture of alcohol and nitromethane, wherein the alcohol is ethanol, and the weight ratio of the nitromethane to the ethanol is 0.5 . 8 to 2 . 5 propellant for turbine-driven gas generators. with added methylamine,
A propellant for a gas generator for driving a turbine according to any one of claims 1 to 3. A method for producing a propellant for a gas generator for driving a turbine according to any one of claims 1 to 4,
a weighing step of respectively weighing the alcohol and the nitromethane;
a mixing step of mixing the alcohol and the nitromethane weighed in the weighing step;
A method for producing a propellant for a gas generator for driving a turbine comprising:

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