JP4840988B2 - Ram Rocket - Google Patents

Description

  The present invention relates to a ram rocket (also referred to as a ram jet) that generates a flying thrust by burning an air-fuel mixture produced by injecting fuel into air introduced from outside in a combustion chamber.

  Conventionally, examples of this type of ram rocket include a ducted rocket and a liquid ram rocket. The former ducted rocket includes a combustion chamber having air intakes at a plurality of positions in the circumferential direction, a solid fuel filled in a gas generator located in front of the combustion chamber, and a ram rocket disposed at the rear end of the combustion chamber. In this ducted rocket, fuel rich combustible gas generated by primary combustion of solid fuel is supplied into the combustion chamber, and a mixture of this combustible gas and air introduced from the air intake is secondary. The combustion gas generated by the combustion is accelerated from the ram rocket nozzle and ejected to generate thrust.

The latter liquid ram rocket is obtained by filling a tank located in front of the combustion chamber with liquid fuel instead of the solid fuel of the ducted rocket. In this liquid ram rocket, the liquid fuel is introduced into the air introduced from the air intake. The liquid fuel droplets and air mixture are burned in the combustion chamber, and the combustion gas generated thereby is accelerated from the ram rocket nozzle and ejected to generate thrust. It has become a thing.
Japanese Patent Laid-Open No. 4-148049 JP-A-5-086980 Japanese Patent Laid-Open No. 4-214953

  However, in the above-mentioned ram rocket, the former ducted rocket has a problem that the specific thrust is low because it is necessary to add an oxidant to the fuel for the purpose of primary combustion of the solid fuel. It was.

  In addition, although the latter liquid ram rocket can obtain a high specific thrust, it cannot be said to have good combustion stability because liquid fuel at room temperature is mixed with air and combusted, and is not as good as a ducted rocket. There is a problem that problems such as ignition and blow-off are likely to occur, and it has been a conventional problem to solve these problems.

  The present invention has been made paying attention to the above-described conventional problems, and achieves high specific thrust equivalent to or close to that of a liquid ram rocket after realizing improvement in ignitability and combustion stability in the combustion chamber. The aim is to provide a ram rocket that is possible.

  A ram rocket according to the present invention is a ram rocket provided with a combustion chamber that mixes and burns fuel and air introduced from the outside, and is equipped with a solid fuel and a liquid fuel. A solid fuel supply means for supplying a combustible gas into the combustion chamber and mixing it with externally introduced air to perform secondary combustion, and supplying liquid fuel droplets into the combustion chamber in which secondary combustion of the solid fuel has started Liquid fuel supply means for mixing with externally introduced air and burning the fuel rich combustible gas supplied to the combustion chamber via the solid fuel supply means, and liquid fuel supply means It is characterized in that the flow rate ratio with the liquid fuel droplets supplied to the combustion chamber is 1: 3.

  In the ram rocket according to the present invention, the fuel-rich combustible gas generated by the primary combustion of the solid fuel is high temperature. First, if the high temperature combustible gas is supplied into the combustion chamber by the solid fuel supply means, the ignitability is improved. In addition, since the supply of the high-temperature combustible gas generated by the primary combustion of the solid fuel into the combustion chamber is continued, the combustion stability is also improved.

  If the liquid fuel supply means supplies the liquid fuel droplets to the combustion chamber in which the secondary combustion of the solid fuel has started in this way, the ratio of the oxidant to the entire fuel is small. As a result, a high specific force equivalent to or close to that of a liquid ram rocket can be obtained.

  According to the ram rocket of the present invention, since it has the above-described configuration, it is possible to improve both ignitability and combustion stability in the combustion chamber, and obtain a high specific thrust equivalent to or close to that of a liquid ram rocket. It has a very good effect that it is possible.

  In the ram rocket of the present invention, a configuration provided with a control valve for controlling the supply amount of liquid fuel droplets may be adopted. When this configuration is adopted, the liquid fuel according to the flight conditions of the ram rocket The supply amount can be controlled.

  In the ram rocket of the present invention, the primary combustion pressure of the solid fuel may be used as the pressurized pressure for supplying the liquid fuel into the combustion chamber. In this case, a pumping means such as a turbo pump is required. It will not be.

  In the hybrid type ram rocket of the present invention, the primary combustion pressure is changed, and the fuel flow rate is changed in accordance with the combustion speed characteristic of the solid fuel. Therefore, the flow rate control by the control valve for controlling the liquid fuel droplet supply amount In order to increase the range, a solid fuel with high pressure sensitivity, that is, a large pressure index is required.

  In this case, in the solid fuel having a pressure index of 0.7 capable of taking a large flow rate control range, the flow rate control range is obtained as follows.

First, the fuel flow rate mf of the hybrid type ram rocket is as follows: solid fuel density ρ _g , combustion area A _b , and combustion speed r
mf = ρ _g A _b r (1)
In this case, the combustion rate r is represented by a case where the combustion rate constant is a, the primary combustion pressure is P ₁ , and the pressure index is n (= 0.7),
r = a (P ₁ ) ⁿ (2)
Therefore, if formula 2 is substituted into formula 1,
mf = ρ _g A _b a (P ₁ ) ⁿ (3)
It becomes.

Here, the primary combustion control range of the pressure _{P 1} and 2MPa～12MPa, mf the flow rate of the solid fuel when the primary combustion pressure _{P 1} is 12 MPa (flow rate of the fuel-rich combustible gas generated by the primary combustion of the solid fuel) _{Assuming 0} , the fuel flow rate mf for any primary combustion pressure P ₁ is given by equation (3):
_{mf = mf 0 (P 1/} 12) n (4)
It becomes. Figure 3 shows the relationship between the primary combustion pressure _{P 1} and mf / mf _0.

In FIG. 3, the value of mf / mf ₀ when the primary combustion pressure P ₁ is 2 MPa is 0.285, and in a ram rocket using a solid fuel with a pressure index of 0.7, the primary combustion pressure control range 2 MPa to At 12 MPa, the fuel flow rate can be changed to about 3.5 times (1 / 0.285).

  In the hybrid type ram rocket of the present invention, the target was set to a fuel flow rate control range in which the fuel flow rate can be changed seven times, and the burning rate characteristics of the solid fuel necessary for this were studied.

First, the primary combustion pressure P ₁ is the ratio of the flow rate of the solid fuel and liquid fuel in the case of 12 MPa 1: 1, the fuel flow rate when blocking the liquid fuel is halved, the primary combustion pressure P ₁ is 2MPa In order for the fuel flow rate to become 1/7 when it changes, the pressure index should be 0.7 from equation (4).

Then, calculating the change in fuel flow to the primary combustion pressure P ₁ at this time. Assuming that the flow coefficient is C _D , the liquid fuel injection hole total area is A _i , the liquid fuel density is ρ _l , the secondary combustion chamber pressure is P ₂ , and the liquid fuel tank is pressurized by the primary combustion pressure P ₁ If the tank pressure is assumed to be equal to the primary combustion pressure P _1, the liquid fuel ml of
ml = C _D A _i (2ρ _l (P ₁ −P ₂ )) ^1/2 (5)
It becomes.

Here, the primary combustion pressure _{P 1} is a flow ml ₀ of the liquid fuel at the time of 12 MPa, secure the flow coefficient _{C D,} when constant at 0.5MPa for simplicity the secondary combustion pressure _{P 2,} any primary combustion liquid fuel flow ml at a pressure P ₁ is determined from the following equation.
ml = ml ₀ ((P ₁ −0.5) / (12−0.5)) ^1/2 (6)

From this relational expression (6), the relationship between the primary combustion pressure P ₁ and (mf + ml) / (mf ₀ + ml ₀ ) is as shown in FIG.

  The minimum value of the fuel flow rate is 0.14 (= 1/7), and it has been found that the fuel flow rate can be changed seven times in the primary combustion pressure control range of 2 MPa to 12 MPa.

The primary combustion pressure P ₁ is the ratio of the flow rate of the solid fuel and liquid fuel in the case of 12 MPa 1: case 2, the fuel flow rate when blocking the liquid fuel becomes 1/3, the primary combustion pressure P ₁ is 2MPa In order for the fuel flow rate to become 1/7 when changed, the pressure index should be 0.47 from equation (4).

The relationship between the primary combustion pressure P ₁ and (mf + ml) / (mf ₀ + ml ₀ ) at this time is as shown in FIG.

The value of (mf + ml) / (mf ₀ + ml ₀ ) when the primary combustion pressure P ₁ is 2 MPa when the liquid fuel valve is opened is 0.38, and the liquid fuel is shut off when the primary combustion pressure P ₁ is 12 MPa. The value of (mf + ml) / (mf ₀ + ml ₀ ) is larger than 0.33 (= 1/3). Therefore, there is a problem that the fuel flow rate of 0.33 to 0.38 cannot be produced. On the other hand, when two liquid fuel shutoff valves are installed and one liquid fuel shutoff valve is opened, half of the liquid fuel injection holes are used. The relationship between the primary combustion pressure P ₁ and (mf + ml) / (mf ₀ + ml ₀ ) when one liquid fuel shut-off valve is opened is also as shown in FIG.

Further, the primary combustion pressure P ₁ is the ratio of the flow rate of the solid fuel and liquid fuel in the case of 12 MPa 1: 3, then the fuel flow rate when blocking the liquid fuel becomes 1/4, the primary combustion pressure P ₁ is 2MPa In order for the fuel flow rate to become 1/7 when it changes, the pressure index should be 0.31 from equation (4).

The relationship between the primary combustion pressure P ₁ and (mf + ml) / (mf ₀ + ml ₀ ) at this time is as shown in FIG. FIG. 6 also shows the relationship between the primary combustion pressure P ₁ and (mf + ml) / (mf ₀ + ml ₀ ) when three liquid fuel shut-off valves are installed.

  As described above, when the ratio of the flow rate of the solid fuel to the liquid fuel is 1: 3, the result that the pressure index is about 0.3 is obtained. Therefore, in the ram rocket of the present invention, the solid fuel supply means is It is desirable that the flow ratio of the fuel-rich combustible gas supplied to the combustion chamber via the liquid fuel droplets supplied to the combustion chamber via the liquid fuel supply means is 1: 3. When this configuration is adopted, it is not necessary to use a solid fuel having a special composition, and an AP-based composite used in a conventional solid rocket propellant can be adopted.

  At this time, it is considered that it is more advantageous to include metal particles in view of the ignition quality and combustion stability of the secondary combustion, and the candidates for the metal particles include B, Mg, which has high combustion heat and is not a rare metal. , Al, Si, P, Ca, Ti, Zr.

  As shown in Table 1, Mg, Al, Si, and Ca are considered to undergo gas phase combustion since the boiling point of the original metal is lower than the boiling point of the oxide, and in particular, P has a very low boiling point itself. It is thought that it vaporizes in the primary combustion chamber. On the other hand, B, Ti, and Zr are considered to undergo surface combustion because the boiling point of the original metal is higher than the boiling point of the oxide.

  In order to improve the ignitability and combustion stability of secondary combustion, it is advantageous to disperse high-temperature metal particles in the combustion field because it is advantageous because the ignition source can be present closely. It is considered that B, Ti, and Zr that maintain the particle state and burn on the particle surface are more suitable as a solid fuel composition than Mg, Al, Si, and Ca.

  Therefore, for B, Ti, and Zr, the theoretical Isp performance of the hybrid type ram rocket was calculated using the chemical equilibrium calculation program CEA400 developed by NASA, and the performance was examined. The studied composition is shown in Table 2.

  The mainstream binder of recent composite propellants is HTPB, but if boron particles are mixed with this HTPB, it cannot be foamed and kneaded, so in the theoretical calculation, CTPB was unified for comparison. The content of metal particles and the content of ammonium perchlorate (AP) were assumed to be 30 wt% metal and 40 wt% AP that could contain the most metal particles in the past results. As for the liquid fuel, JP-4, which is used as a fuel for aircraft engines, was studied, and the flight conditions were assumed to be altitude 0 km and speed Mach 2.

When the optimum aperture ratio on the ground in the rocket performance calculation result of CEA400 is Isp _mix , Isp as a hybrid type ram rocket is obtained as follows.
First, if the thrust is Fn, the fuel flow rate is mf, and the gravitational acceleration is g,
Isp = Fn / mf · g (7)
In this case, assuming that the thrust generated at the nozzle is Fg, the air flow rate is ma, and the inflow air flow velocity is v,
Fn = Fg−ma · v (8)
Is represented by
Isp _mix = Fg / (mf + ma) · g (9)
Here, if Expressions 8 and 9 are substituted into Expression 7,
Isp = (Isp _mix (mf + ma) · g−ma · v) / mf · g
= Isp _mix (1 + A / F) − (A / F · v) / g (10)
It becomes. However, ma / mf = A / F

  Therefore, FIG. 7 shows the calculation result of the theoretical performance of Isp with respect to A / F when there is no liquid fuel but only solid fuel. As can be seen from FIG. 7, the B-containing fuel having the highest combustion heat per mass has the highest Isp value, and then the Isp value is higher in the order of Ti-containing fuel and Zr-containing fuel. At A / F = 15, there is a difference of Isp200s or more between the B-containing fuel and the Ti, Zr-containing fuel.

  FIG. 8 and FIG. 8 show the calculation results of the theoretical performance of Isp with respect to A / F when the ratio of the flow rate of the solid fuel and the liquid fuel is 1: 1 and when the ratio of the flow rate of the solid fuel and the liquid fuel is 1: 3. 9 respectively.

  8 and 9, when the ratio of the flow rates of the solid fuel and the liquid fuel is 1: 1, the difference in Isp value between the B-containing fuel and the Ti, Zr-containing fuel is 100 s or less at A / F = 15. When the ratio of the flow rates of the solid fuel and the liquid fuel is 1: 3, it can be seen that there is almost no difference in the Isp value between the B-containing fuel and the Ti, Zr-containing fuel at A / F = 15.

  Therefore, in the hybrid type ram rocket, in terms of theoretical performance, any of the B-containing fuel, Ti-containing fuel and Zr-containing fuel is almost equivalent, so which composition is used depends on the secondary combustion ignitability, It is necessary to determine from the secondary combustion stability and the damage status of the secondary combustion chamber insulation.

  EXAMPLES Hereinafter, although an Example demonstrates this invention still in detail, this invention is not limited to a following example.

  1 and 2 show a ram rocket according to an embodiment of the present invention.

  As shown in FIGS. 1 and 2, the ram rocket 1 has a combustion chamber 3 loaded with a booster propellant 2, and a booster nozzle 4 and a ram rocket nozzle 5 are provided at the rear end of the combustion chamber 3. A booster nozzle 4 is provided with a booster propellant ignition device 7 concentrically provided via a separation joint 6.

  Air intakes 8 disposed at a plurality of locations in the circumferential direction are provided on the outer periphery on the front end side of the combustion chamber 3, and the port cover 9 that closes the air intake 8 is released by the port cover release mechanism 10. It has become.

  In this case, a gas generation chamber 17 filled with a solid fuel 16 is provided in front of the combustion chamber 3. The gas generation chamber 17 is provided with a solid fuel ignition device 18 for igniting the solid fuel 16, and a gas injection nozzle 19 as a solid fuel supply means provided integrally with the gas generation chamber 17. The gas injection nozzle 19 penetrates the partition wall 20 and protrudes into the combustion chamber 3.

  Further, in front of the combustion chamber 3, a fuel tank 12 filled with the liquid fuel 11, a flow rate control valve 13 for controlling the flow rate of the liquid fuel 11, and liquid fuel supply means for guiding the liquid fuel 11 to the air intake 8. The liquid fuel 11 in the fuel tank 12 is supplied into the combustion chamber 3 (air intake 8) via the conduit 14 by the combustion pressure generated by the primary combustion of the solid fuel 16. It looks like this.

  In this embodiment, the fuel-rich combustible gas generated by primary combustion of the solid fuel 16 supplied to the combustion chamber 3 through the gas injection nozzle 19 and the liquid fuel supplied to the combustion chamber 3 through the conduit 14 are used. The flow rate ratio with respect to the droplets is set to 1: 3, and the solid fuel 16 is made of an AP-based composite containing any metal particle of B, Ti, and Zr.

  In the ram rocket 1 having such a configuration, the booster propellant 2 is ignited by the booster propellant ignition device 7 to burn the booster propellant 2 in the combustion chamber 3, and this combustion gas is injected from the booster nozzle 4. The vehicle starts and accelerates until it reaches the set Mach number necessary for operation by ram pressure.

  Subsequently, when the speed reaches the set Mach number and the combustion of the booster propellant 2 is completed, the booster nozzle 4 is separated at the separation joint 6 as shown in FIG. Thus, the port cover 9 of the air intake 8 is removed.

  In accordance with this, first, the solid fuel 16 is ignited by the solid fuel ignition device 18, and a fuel-rich combustible gas generated by primary combustion of the solid fuel 16 is passed through the gas injection nozzle 19 to the combustion chamber 3. The combustion gas for thrust generation generated by the combustion of the air-fuel mixture in the combustion chamber 3 is mixed with the external air supplied from the air intake 8 and mixed with the air from the outside. Accelerates from and squirts to obtain thrust.

  At this time, since the fuel-rich combustible gas generated by primary combustion of the solid fuel 16 is high temperature, the ignitability in the combustion chamber 3 is improved.

  Then, droplets of the liquid fuel 11 are supplied through the conduit 14 into the combustion chamber 3 (air intake 8) where the secondary combustion of the solid fuel 16 is started as described above. The mixture of the liquid droplets and the external air introduced from the air intake 8 is ignited and burned by the combustion gas generated by the secondary combustion of the solid fuel 16, and the combustion gas is accelerated from the ram rocket nozzle 5. Spout to gain more thrust.

  That is, when the droplets of the liquid fuel 11 are supplied into the combustion chamber 3 where the secondary combustion of the solid fuel 16 has started via the conduit 14 and mixed with the introduced air from the outside and burned. As a result, the ratio of the oxidizer to the whole fuel is reduced, and as a result, a high specific thrust equivalent to or close to that of a liquid ram rocket can be obtained.

  In addition, the supply of the high-temperature combustible gas generated by the primary combustion of the solid fuel 16 to the combustion chamber 3 is continuously performed through the gas injection nozzle 19, so that the air and the droplets of the liquid fuel 11 are mixed. Even after ignition of the air-fuel mixture, the air-fuel mixture is sustained not only by its own combustion energy but also by the combustion gas ejected from the gas injection nozzle 19, so that high altitude flight can be achieved. Even when the gas flow rate in the combustion chamber 3 is large and the pressure is low as in the case of the design standard, the combustion of the air-fuel mixture is favorably sustained without causing blowout.

  In the ram rocket 1 described above, since the flow rate control valve 13 for controlling the supply amount of the liquid fuel 11 droplets is provided, it is possible to control the supply amount of the liquid fuel 11 according to the flight conditions of the ram rocket 1.

  Further, in the ram rocket 1 described above, the liquid fuel 11 in the fuel tank 12 is supplied into the combustion chamber 3 (air intake 8) via the conduit 14 by the combustion pressure generated by the primary combustion of the solid fuel 16. Therefore, the liquid fuel 11 is pressed at a high pressure, and the liquid fuel 11 in the fuel tank 12 is pumped to the air intake 8 without requiring a turbo pump.

  Further, in the ram rocket 1 described above, the flow rate of the fuel-rich combustible gas supplied to the combustion chamber 3 via the gas injection nozzle 19 and the liquid fuel droplets supplied to the combustion chamber 3 via the conduit 14. Since the ratio is 1: 3, it is not necessary to use a fuel with a special composition as the solid fuel 16 (for example, a high-energy binder fuel such as GAP), and the AP system used in the propellant of a conventional solid rocket Composites can be used.

  Furthermore, in the ram rocket 1 described above, as the solid fuel 16, an AP-based composite containing one of B, Ti, and Zr, which has high combustion heat and is not a rare metal, can be used. B-containing fuel, Ti-containing fuel and Zr-containing fuel with little difference in theoretical performance can be selected as appropriate according to secondary combustion ignitability, secondary combustion stability, damage status of secondary combustion chamber insulation, etc. It will be.

It is sectional explanatory drawing which shows one Example of the ram rocket concerning this invention. Example 1 FIG. 2 is a cross-sectional explanatory view showing a state of the booster propellant combustion end stage of the ram rocket of FIG. 1. Example 1 It is a graph showing the relationship between the primary combustion pressure _{P 1} and mf / mf _0. The ratio of the flow rate of the solid fuel and liquid fuel when the primary combustion pressure _{P 1} is 12MPa is 1: is a graph showing the relationship between 1 primary combustion pressure _{P 1} and in the case of _{(mf + ml) / (mf} 0 + ml 0). The ratio of the flow rate of the solid fuel and liquid fuel when the primary combustion pressure _{P 1} is 12MPa is 1: is a graph showing the relationship between the primary combustion pressure _{P 1} and in the case of _{2 (mf + ml) / (} mf 0 + ml 0). Primary combustion pressure _{P 1} is the ratio of the flow rate of the solid fuel and liquid fuel in the case of 12 MPa 1: is a graph showing the relationship between the primary in the case of the tertiary combustion pressure _{P 1} and _{(mf + ml) / (mf} 0 + ml 0). It is a graph which shows the calculation result of the theoretical performance of Isp with respect to A / F when there is no liquid fuel but only solid fuel. It is a graph which shows the calculation result of the theoretical performance of Isp with respect to A / F in case the ratio of the flow volume of a solid fuel and a liquid fuel is 1: 1. It is a graph which shows the calculation result of the theoretical performance of Isp with respect to A / F in case the ratio of the flow volume of a solid fuel and a liquid fuel is 1: 3.

Explanation of symbols

DESCRIPTION OF SYMBOLS 1 Ram rocket 3 Combustion chamber 11 Liquid fuel 13 Flow control valve 14 Conduit (liquid fuel supply means)
16 Solid fuel 19 Gas injection nozzle (solid fuel supply means)

Claims (3)

In a ram rocket with a combustion chamber that mixes and burns fuel and air introduced from the outside,
A solid fuel supply means for mounting a solid fuel and a liquid fuel, supplying a fuel-rich combustible gas generated by primary combustion of the solid fuel into the combustion chamber, mixing it with externally introduced air, and performing secondary combustion; and solid fuel Liquid fuel supply means for supplying liquid fuel droplets into the combustion chamber in which the secondary combustion is started and mixing with externally introduced air to burn is provided ,
The flow ratio between the fuel-rich combustible gas supplied to the combustion chamber via the solid fuel supply means and the liquid fuel droplets supplied to the combustion chamber via the liquid fuel supply means is 1: 3. Ram rocket characterized by.   The ram rocket according to claim 1, further comprising a control valve for controlling a supply amount of liquid fuel droplets.   The ram rocket according to claim 1 or 2, wherein the primary combustion pressure of the solid fuel can be used as a pressurized pressure for supplying the liquid fuel into the combustion chamber.

Images