JP2006152917A - Nozzleless roket - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a nozzleless rocket suppressing unstable combustion or lowering of combustion pressure of a propellant to increase the range of a flying body and improve a flying performance such as speeding up. <P>SOLUTION: In the nozzleless rocket, an ignition device 2 is disposed to a front part of a cylindrical combustion case 1, a first propellant 4 and a second propellant 5 fill, in layers, a combustion chamber 3 in the combustion case 1, and an inner hole 7 extending in a direction of center axis 6 of the combustion case 1 is formed at the central part of the first propellant. The sum of combustion times from an inner surface 4b of the first propellant 4 forming the inner hole 7 to an inner wall face 1a of the combustion case 1 in a radial direction for each position in the center axis 6 direction has the same time length. The nozzleless rocket is suitably used as a booster applicable to a ducted rocket engine. <P>COPYRIGHT: (C)2006,JPO&NCIPI

Description

  The present invention relates to a nozzleless rocket and a nozzleless rocket for a ducted rocket engine using the same as a booster (auxiliary propulsion device).

  Conventionally, nozzleless rockets that do not have nozzles, which are components of rockets, have been proposed from the viewpoint of improving reliability based on simplification of the structure by reducing the number of parts. As such a nozzleless rocket, what is shown in FIG. 12 is known (for example, refer nonpatent literature 1).

  That is, as shown in FIG. 12, an ignition device 34 including an ignition powder 32 and a squib 33 is attached to the front end of the cylindrical combustion case 31. The combustion chamber 35 in the combustion case 31 is filled with a propellant 36, and a circular hole 38 is formed in the center of the propellant 36, that is, in the direction of the central axis 37 of the combustion case 31. In this nozzleless rocket, the igniter 32 is ignited by the squib 33, and the inner surface 36a of the propellant 36 is ignited and combusted by the ignited gas. Thereafter, the generated combustion gas flows through the circular hole 38 of the propellant 36 while increasing the flow velocity and flow rate, and thrust is generated by jetting the combustion gas at a high speed rearward.

However, such a nozzleless rocket has the following problems.
As shown in FIG. 13, the flow rate of the combustion gas flowing through the circular hole 38 of the propellant 36 increases from the front end toward the rear end, and the pressure decreases. On the other hand, the combustion speed of the propellant 36 is affected by the pressure and flow velocity of the combustion gas on the combustion surface, and as shown in FIG. 14 (a), the combustion speed increases as the pressure by the combustion gas increases. When a combustion gas flow is present on the surface as shown in FIG. 14 (b), the combustion rate further increases as the flow velocity of the combustion gas increases.

  Under such circumstances, when the degree of increase in the combustion speed of the propellant 36 due to the increase in the flow velocity of the combustion gas flowing over the surface is larger than the degree of decrease in the combustion speed due to decrease in the pressure of the combustion gas, the central axis The burning speed is not uniform in the 37 direction, the front part burns slowly, and the rear part burns fast. On the contrary, when the degree of increase in the combustion speed of the propellant 36 due to the increase in the flow velocity of the combustion gas flowing on the surface is smaller than the degree of decrease in the combustion speed due to decrease in the pressure of the combustion gas, the combustion speed in the direction of the central axis 37. Is not uniform, the front part burns fast and the rear part burns slowly.

  As a result, for example, as shown in FIG. 15, the shape of the propellant 36 changes from the time t1, t2, t3, t4, and t5 after ignition, and the propulsion that forms a partial circular hole 38 of the propellant 36. After the time point t3 when the inner surface 36a of the medicine 36 reaches the inner peripheral surface 31a of the combustion case 31, the reduction in the combustion area, and hence the combustion pressure, becomes significant. Then, unstable combustion as shown in FIGS. 16 (a) and 16 (b) peculiar to the solid rocket motor in the low pressure region occurs, or the combustion is interrupted and the propellant 36 remains. As a result, there have been problems that the speed of the flying object is not sufficient or that the component parts are damaged by vibration.

  Furthermore, in recent years, in a ducted rocket engine, a nozzleless rocket that does not need to separate the booster rocket nozzle after accelerating the flying object with a booster rocket is proposed as a booster rocket for a ducted rocket engine as shown in FIG. (For example, see Non-Patent Document 2).

  That is, as shown in FIG. 17, a gas generator 43 is disposed at the front of the ducted rocket engine 41, and a nozzleless rocket 45 having a combustion chamber 35 (secondary combustion chamber) is disposed at the rear. Between them, an injection hole 46 communicated between the two and a flow rate control device 47 for controlling the flow rate of the combustion gas passing therethrough are provided. An air intake duct 48 is provided at the side of the ducted rocket engine 41 so as to introduce air into the combustion chamber 35. A port cover 49 is provided at the inlet of the air intake duct 48 to the combustion chamber 35, and the air is taken into the combustion chamber 35 by being destroyed at a predetermined time. A heat insulating material (not shown) is disposed on the inner peripheral portion of the combustion case 31, and the inside thereof becomes the combustion chamber 35, and the combustion chamber 35 is filled with the propellant 36. A ram nozzle 50 for a ducted rocket engine is provided at the rear end of the ducted rocket engine 41.

The propellant 36 accommodated in the combustion chamber 35 is ignited and burned by the operation of the ignition device 34, and the flying object is accelerated to a predetermined speed by the generated thrust. Thereafter, the port cover 49 is released, and the air is compressed through the air intake duct 48 and taken into the combustion chamber 35. At the same time, the operation of the gas generator 43 disposed in front of the combustion chamber 35 is started, and the generated fuel gas is injected into the combustion chamber 35 through the injection hole 46, and mixed with the air and ignited to start combustion. To generate thrust. Thereafter, the flying object generates thrust while the amount of gas generated from the gas generator 43 is optimally controlled by the flow control device 47 in accordance with the amount of air taken in due to the change in flight altitude and speed, and the flight continues.
Alon Gany and Israel Aharon, "Internal Ballistics Considerations of Nozzleless Roket Motors", JOURNAL OF PROPULSION AND POWER, Vol.15, No.6 November-December 1999 URL: http://www.bayernchemie-protac.com/en/products/airbreathing / meteor BAYERN-CHEMIE PROTAC website (searched July 23, 2004)

  However, in the ducted rocket engine 41, the combustion speed in the direction of the central axis 37 is not constant as described above for the propellant of the nozzleless rocket applied to the booster. Unstable combustion as shown in FIG. 16 occurs. Due to such unstable combustion, the release time of the port cover 49 for the operation of the ducted rocket engine is delayed, the time for no-thrust flight becomes longer, the speed decreases due to the air resistance during that time, and the predetermined range cannot be obtained. There was a problem. Further, even if unstable combustion does not occur in the late stage of combustion due to the action of the ram nozzle, the combustion pressure is lowered and the performance is lowered. As a result, it is necessary to increase the propellant amount in order to obtain a necessary speed. Therefore, it is necessary to lengthen the combustion chamber 35. As a result, the weight of the ducted rocket engine 41 is increased, and the flight performance such as a reduction in the speed of the flying object is deteriorated.

  Therefore, an object of the present invention is to provide a nozzleless rocket capable of suppressing the unstable combustion of the propellant or the decrease in the combustion pressure, and improving the flight performance such as extending the range of the flying object and increasing the speed. It is to provide.

  In order to achieve the above object, the nozzleless rocket according to the first aspect of the present invention is provided with an ignition device at the front of a cylindrical combustion case, and a propellant is placed in the combustion chamber in the combustion case. In a nozzleless rocket which is filled and has an inner hole extending in the central axis direction of the combustion case at the center of the propellant, at least two types of propellants are radially spread over the whole area or a partial area in the central axis direction. It is configured so that the sum of the combustion time until reaching the inner wall surface of the combustion case in the radial direction from the inner surface of the propellant which is filled in layers and forms an inner hole at each position in the central axis direction is the same. It is a feature.

  In the nozzleless rocket of the second invention, in the first invention, the radial thickness of each propellant constituting at least two types of propellants is the composition of the propellant and the diameter of the combustion chamber filled with the propellant. And the length, and the cross-sectional area and length of the inner hole are set.

  The nozzleless rocket of the third invention is provided with a gas generator for generating fuel gas at the front part, the nozzleless rocket of the first or second invention at the rear part, and air on the side part. An intake duct is provided, which is used as a booster applied to a ducted rocket engine in which fuel gas from a gas generator and air from an air intake duct are mixed and burned.

According to the present invention, the following effects can be exhibited.
In the nozzleless rocket according to the first aspect of the present invention, at least two kinds of propellants are filled in layers in the radial direction over the entire region or a partial region in the central axis direction, and propulsions that form inner holes at respective positions in the central axis direction The sum of the combustion time from the inner surface of the medicine to the inner wall surface of the combustion case in the radial direction is the same. Therefore, the time until the inner surface of the propellant that forms the inner hole during combustion of the propellant reaches the inner wall surface of the combustion case becomes uneven, resulting in a decrease in the combustion area, that is, a decrease due to a decrease in combustion pressure. The propellant burns stably without causing stable combustion, and can suppress a decrease in combustion pressure. As a result, it is possible to improve the flight performance such as extending the range of the flying object and increasing the speed.

  Further, in the nozzleless rocket of the second invention, in order to make the sum of the combustion times of the propellants described above, the thickness in the radial direction of each propellant constituting at least two types of propellants is It is set based on the composition, the diameter and length of the combustion chamber filled with the propellant, and the cross-sectional area and length of the inner hole. Thus, since the radial thickness of the propellant is set based on more specific conditions, the effect of the first invention can be sufficiently exerted.

  Furthermore, in the nozzleless rocket of the third invention, the nozzleless rocket of the first or second invention is used as a booster applied to a ducted rocket engine. Therefore, the effect of the nozzleless rocket of the first or second invention can be exhibited as a booster of a ducted rocket engine.

(First embodiment)
Hereinafter, a first embodiment of the present invention will be described in detail with reference to the drawings.
1-7 is a figure for demonstrating 1st Embodiment of this invention. FIG. 1 is a cross-sectional view showing an example of a nozzleless rocket, FIGS. 2A and 2B are graphs showing the combustion rate characteristics of the propellant used, and FIGS. 3A, 3B and 3C are FIG. 3 is a partial enlarged cross-sectional view for explaining the thickness setting of the propellant used. 4 (a) to 4 (e) are cross-sectional views showing a molding process of a propellant used for a nozzleless rocket, FIG. 5 is a graph showing a pressure pattern of the nozzleless rocket, and FIG. 6 is a propellant used for the nozzleless rocket. FIG. 7 is a graph showing the combustion rate characteristics of the propellant in FIG. 6, and FIG. 8 is a graph showing the pressure pattern of the nozzleless rocket in FIG.

  In the nozzleless rocket of the present invention, an ignition device is provided at the front of the combustion case, and the combustion chamber in the combustion case is filled with propellant. The propellant has an inner hole in the central axis direction of the combustion case, and is configured such that at least two types of propellants are filled in layers in the radial direction over the entire region or a partial region in the central axis direction. The propellant is configured such that the sum of combustion times from the inner surface of the propellant forming the initial inner hole at each position in the central axis direction to the inner wall surface of the combustion case in the radial direction is the same. Such a nozzleless rocket will be specifically described below.

  As shown in FIG. 1, an ignition device 2 is attached to the front of an iron combustion case 1 having a covered cylindrical shape, and two kinds of propellants, that is, a first propellant 4 are placed in the combustion chamber 3 therein. And the second propellant 5 are filled in layers in the radial direction. An inner hole 7 extending in the direction of the central axis 6 of the combustion case 1 is provided at the center of the first propellant 4. In this example, two types of propellants, the first propellant 4 and the second propellant 5, are used, but three or more types of propellants can also be used. Considering the productivity of the propellant described later, it is preferable to use two types of propellants as in this example.

  The first propellant 4 and the second propellant 5 have different compositions and therefore different combustion rates. In this example, the second propellant 5 is faster than the combustion rate of the first propellant 4. The radial thickness 4a of the first propellant 4 is increased from the front portion of the combustion chamber 3 toward the rear portion, and conversely, the second propellant 5 is decreased from the front portion of the combustion chamber 3 toward the rear portion. It is configured. However, the inner hole 7 is formed to have the same diameter from the front part to the rear part.

  The feature of the nozzleless rocket is that from any inner surface 4b of the first propellant 4 that forms the initial inner hole 7 to the inner wall surface 1a of the combustion case 1 in the radial direction at any position in the central axis 6 direction of the propellant. The sum of the combustion times, that is, the sum of the combustion times of the first propellant 4 and the second propellant 5 is the same. Here, in the case where a heat insulating material is provided on the inner wall surface 1 a of the combustion case 1, the combustion time is the heat insulating material in the radial direction from the inner surface 4 b of the first propellant 4 that forms the initial inner hole 7. It will be the combustion time to the inner surface of.

  In this example, the second propellant 5 having a higher burning rate than that of the first propellant 4 is thickly filled in a region where the burning rate of the first propellant 4 is slow, and the second propellant 4 is thinly packed in a region where the burning rate of the first propellant 4 is high. It is formed in a tapered shape. In this case, a configuration in which the first propellant 4 is not filled on the rear side in the direction of the central axis 6 can also be adopted. However, on the contrary, when the second propellant 5 having a slower combustion rate than the first propellant 4 is used, the second propellant 5 is filled thickly in the region where the combustion rate of the first propellant 4 is fast, A thin filling or non-filling configuration can also be adopted. With such a configuration, the sum of the burning times in the radial direction of both propellants can be made the same at any position in the direction of the central axis 6.

  In this example, as the first propellant 4 and the second propellant 5, a composite propellant (solid propellant) in which an oxidant and a fuel are mixed is used. That is, a composite propellant whose oxidizer is a perchlorate such as ammonium perchlorate (AP) and whose binder (fuel and binder) is a hydroxyl group-containing polymer such as terminal hydroxyl group-containing polybutadiene (HTPB) as a main component. Used. A propellant composed of a double base propellant other than the composite propellant can be used as long as the condition that the sum of the combustion times is the same can be satisfied and there is no problem in manufacturability.

  In order to set the radial thickness 4a of the first propellant 4 and the radial thickness 5a of the second propellant 5, for example, the following method can be used. Conventionally, to calculate the combustion performance of a rocket propellant, a calculation program that uses the pressure of the combustion gas flowing on the surface of the propellant, the combustion velocity characteristics of the propellant relative to the flow velocity, or the shape of the propellant as input data has been used. Changes in the inner shape of the drug are calculated.

  Specifically, for example, the sum of the values obtained by dividing the thickness of each propellant by the combustion speed is made constant so that the sum of the combustion times is the same. That is, the sum of the value obtained by dividing the thickness 4a of the first propellant 4 by the combustion rate and the value obtained by dividing the thickness 5a of the second propellant 5 by the combustion rate is made constant. The burning rate of the first propellant 4 and the burning rate of the second propellant 5 are determined based on the composition of the first propellant 4 and the composition of the second propellant 5 separately. Therefore, it can be easily configured so that the sum of the combustion times is the same in the direction of the central axis 6.

  When the combustion performance calculation program is used, the propellant combustion determined by the composition of the first propellant 4 and the second propellant 5 as shown in FIGS. 2 (a) and 2 (b) as input data. A combustion rate characteristic with respect to pressure and a combustion rate characteristic with respect to the flow velocity of the combustion gas flowing on the combustion surface are input. Further, as input data, the shape of the first propellant 4 (outer diameter D1, inner diameter D2, length L) and the shape of the second propellant 5 shown in FIG. 3A [center axial direction position (X) and its position The inner diameter Dx] of the second propellant 5 is input, and the change in the shape of the propellant can be calculated by the combustion performance calculation program.

  FIGS. 3B and 3C are examples of calculation results when the shape of the propellant is changed. As shown by the two-dot chain line in FIG. 3B, the inner shape of the initial second propellant 5 is shown. Is changed to shape A, shape B, or shape C. In those cases, as shown by a two-dot chain line in FIG. 3 (c), the shape of the inner hole of the propellant at the time when the inner surface 4b of the first propellant 4 first reaches the inner wall surface 1a of the combustion case 1, It is calculated for each of the shape A, shape B, or shape C. By performing calculation by changing the initial shape in various ways in this manner, it is possible to set the initial propellant thickness at which the combustion time of the propellant becomes the same at each position in the direction of the central axis 6 by trial and error.

  The propellant having the shape set in this way shows the combustion pressure pattern shown by the solid line in FIG. 5, and the propellant is stable without dropping to a low pressure level that causes unstable combustion over the entire combustion period. To complete combustion. In FIG. 5, the combustion pressure pattern of the conventional nozzleless rocket is shown together with a broken line.

  Next, a method for forming the propellant filled in the nozzleless rocket will be described with reference to FIG. First, as shown in FIG. 4A, a columnar first mold 8 is mounted on a central axis 6 of a cylindrical combustion case 1 with a fixture 9, and a gap 10 between the first mold 8 and the combustion case 1 is attached. The second slurry 11 serving as the second propellant 5 is poured into the container, and is kept for several days at a temperature necessary for curing to be cured. Thereafter, as shown in FIG. 4 (b), the first mold 8 is released, and then, as shown in FIG. 4 (c), a rotary shaft body 13 having a cutting tool 12 at its tip is formed into a cylindrical shape. 2 Insert into the internal space of the propellant 5. Then, the rotary shaft body 13 is rotated and finished by cutting the inner surface 5b of the second propellant 5 into a predetermined shape (dimension) with the cutting tool 12 at the tip. In this case, surface treatment can be applied to the inner surface 5b of the second propellant 5 as necessary.

  Thereafter, the second mold 14 shown in FIG. 4 (d) is mounted on the central axis 6 of the combustion case 1 by the fixture 9, and the first propellant 4 and the second propellant 4 are placed in the space 15 between the second mold 14 and the second propellant 5. The first slurry 16 is poured and cured. Then, the 2nd casting_mold | template 4 and the 2nd propellant 5 of the shape shown in FIG.4 (e) are laminated-molded by releasing the 2nd casting_mold | template 14 and finishing an end surface. When the shape of the second propellant 5 becomes a set value in the state of FIG. 4B, the inner surface 5b of the second propellant 5 by the cutting tool 12 shown in FIG. Finishing can be omitted.

  Next, a specific example of the propellant manufactured according to the molding process will be described based on the above calculation result. First, as the first propellant 4, a mixture containing 10 to 20% by mass of a terminal hydroxyl group-containing polybutadiene (HTPB) binder and 80 to 90% by mass of ammonium perchlorate (AP) was used. As the 2nd propellant 5, what added 1-10 mass parts of combustion catalysts which are a metal oxide with respect to those HTPB and AP with respect to those 100 mass parts in total was used. As the combustion case 1, a case having a length of 1000 mm and an outer diameter of 100 mm was used. When the initial bore diameter of the first propellant 4 is set to 40 mm, the initial shape of the propellant with the same combustion time for the propellant at each position in the direction of the central axis 6 is set as shown in FIG. The pattern of the burning rate of the 1st propellant 4 and the 2nd propellant 5 was shown in FIG. 7, and the pattern of the combustion pressure using these propellants was shown in FIG. As shown in FIG. 8, the propellant burned stably without dropping to a low pressure level that would cause unstable combustion over the entire combustion period.

  In the example of FIG. 1 described so far, the second propellant 5 is filled so that the thickness becomes zero at the rear end of the nozzleless rocket, but the composition of the propellant, the combustion chamber 3 in which the propellant is filled. Depending on the shape and size, as shown in FIG. 9, the second propellant 5 may be filled in a partial region (front half) in the direction of the central axis 6.

Next, the operation of the nozzleless rocket of the first embodiment will be described.
In the nozzleless rocket shown in FIG. 1, the first propellant 4 is ignited and burned by the ignition device 2, and the combustion gas of the first propellant 4 passes through the inner hole 7 of the first propellant 4 from the front to the rear. To flow. In the conventional nozzleless rocket, as shown in FIG. 13, the flow velocity of the combustion gas increases toward the rear, and the pressure decreases. However, in the first propellant 4 of FIG. 1, the degree of increase in the combustion speed due to the increase in the flow velocity of the combustion gas flowing on the surface becomes larger than the degree of decrease in the combustion speed due to the decrease in the pressure of the combustion gas. The burning speed is not uniform in the direction, the front part is slow, and the rear part burns fast.

  Further, in the portion filled with the second propellant 5 at the front part, after the inner surface 4b of the first propellant 4 forming the inner hole 7 reaches the inner surface 5b of the second propellant 5. The combustion rate increases and the combustion rate becomes faster than the propellant located at the rear. As a result, the arrival time from the inner surface 4b of the first propellant 4 to the inner wall surface 1a of the combustion case 1 at the initial stage at each position in the central axis 6 direction is constant (uniform). As a result, it is possible to avoid a decrease in the combustion area, that is, a low pressure level that causes unstable combustion, and combustion of the propellant is stably performed and required throughout the entire combustion period. Performance is obtained.

The effect exhibited by the above first embodiment will be described below.
In the nozzleless rocket according to the first embodiment, the first propellant 4 and the second propellant 5 are filled in layers in the radial direction over the entire region in the direction of the central axis 6, and the inner holes are formed at respective positions in the direction of the central axis 6. 7 is configured such that the sum of the combustion time from the inner surface 4b of the first propellant 4 forming 7 to the inner wall surface 1a of the combustion case 1 in the radial direction is the same. Accordingly, the combustion area decreases due to nonuniform time until the inner surface 4b of the first propellant 4 forming the inner hole 7 reaches the inner wall surface 1a of the combustion case 1 during the combustion of the propellant. That is, the first propellant 4 and the second propellant 5 are stably burned without causing unstable combustion due to a decrease in the combustion pressure, and a decrease in the combustion pressure can be suppressed. As a result, it is possible to improve the flight performance such as extending the range of the flying object and increasing the speed. In addition, the flying object is not subject to harmful vibrations.

In addition, in order to make the sum of the combustion times of the first propellant 4 and the second propellant 5 described above, the radial thickness of the first propellant 4 and the second propellant 5 is determined by the composition of the propellant. The diameter and length of the combustion chamber 3 filled with the propellant and the cross-sectional area and length of the inner hole 7 are set. Thus, since the thickness of the 1st propellant 4 and the 2nd propellant 5 in the radial direction is set based on more specific conditions, the above-described effects can be sufficiently exerted.
(Second Embodiment)
Next, an example in which the nozzleless rocket of the first embodiment is used as a booster for a ducted rocket engine will be described.

  FIG. 10 is a cross-sectional view showing a ducted rocket engine in which a nozzleless rocket is mounted as a booster for a ducted rocket engine, and differs from the conventional ducted rocket engine shown in FIG. Other configurations are the same. That is, a gas generator 18 is disposed at the front portion of the ducted rocket engine 17, and a nozzleless rocket 20 having the combustion chamber 3 is disposed at the rear portion thereof. An injection hole 21 and a flow rate control device 22 that controls the flow rate of the combustion gas passing through the injection hole 21 are provided. The configuration of the nozzleless rocket 20 is the same as the configuration of FIG. An air intake duct 23 is provided along the axial direction of the ducted rocket engine 17 so as to introduce air into the combustion chamber 3. A port cover 24 is provided at the inlet of the air intake duct 23 to the combustion chamber 3 so that the air is taken into the combustion chamber 3 by being broken at a predetermined time.

  A heat insulating material (not shown) is arranged on the inner peripheral portion of the combustion case 1, and the inside thereof becomes the combustion chamber 3, and the combustion chamber 3 is filled with propellant. An annular recess 25 for attaching a ram nozzle is provided at the rear end of the ducted rocket engine 17 so that a ram nozzle (not shown) can be attached.

Next, the operation of the second embodiment will be described.
The first propellant 4 starts to combust by the operation of the ignition device 2, and the combustion proceeds by the same action as that of the first embodiment in the initial operation. The combustion of the propellant after the inner surface 4b of the first propellant 4 reaches the throat surface of the ram nozzle (not shown) proceeds with the same action as a normal rocket motor in which the combustion gas is accelerated by the ram nozzle. . At that time, the thickness 4a of the first propellant 4 and the thickness 5a of the second propellant 5 are set so that the sum of the combustion times of the first propellant 4 and the second propellant 5 is the same. The operation is stably performed as shown in the pressure pattern of FIG. For this reason, the nozzleless rocket 20 does not cause a large decrease in combustion pressure and unstable combustion as shown in FIG. Thereafter, the port cover 24 is released and the operation of the ducted rocket engine 17 is started.

  Therefore, the flying object equipped with the ducted rocket engine 17 can obtain a predetermined range without causing a speed reduction due to low thrust or no-thrust flight due to unstable combustion. In addition, it is not necessary to extend the combustion chamber 3 to increase the amount of propellant in order to compensate for the performance degradation due to unstable combustion, resulting in a decrease in flight performance due to an increase in the weight of the flying object, which was a problem of the prior art. Nor.

  In the second embodiment, the inner hole 7 of the first propellant 4 is tapered at the rear end, but the cross-sectional shape is the same in the direction of the central axis 6 as in the first embodiment shown in FIG. It can also be. The nozzleless rocket of the present invention can also be applied to a booster of a ramjet engine that uses a liquid fuel tank filled with liquid fuel instead of the gas generator 18.

  As described above, according to the second embodiment, since the combustion time in the radial direction of the propellant at the respective positions in the direction of the central axis 6 is the same, the combustion pressure is lowered to a low level that causes unstable combustion. Without any problem, stable combustion of the propellant can be obtained throughout the entire combustion period. As a result, it is possible to suppress a decrease in flying performance and an increase in weight, and it is possible to increase the speed of the flying object and extend the range. Therefore, the effect of the nozzleless rocket 20 can be effectively exhibited as a booster of the ducted rocket engine 17.

It should be noted that the above-described embodiment can be modified as follows.
-The shape of the boundary part of at least 2 types of propellant can also be comprised so that it may become a curved surface by the cross-sectional quadratic curve etc. in the center axis 6 direction.

It is also possible to change the number of layers of at least two propellants at the front and rear in the direction of the central axis 6, for example, two layers at the front and three layers at the rear.
The inner hole 7 of the first propellant 4 can be tapered so that the diameter of the inner hole 7 increases or decreases toward the rear.

Furthermore, the technical idea grasped from the embodiment will be described below.
The propellant is composed of two types of propellants having different burning rates. The propellant having a low burning rate is formed in a taper shape whose diameter is enlarged toward the rear part of the combustion case. The nozzleless rocket according to any one of claims 1 to 3, wherein the nozzleless rocket is formed in a tapered shape whose diameter is increased toward a front portion. When configured in this way, the sum of the combustion time from the inner surface of the propellant to the inner wall surface of the combustion case in the radial direction at each position in the central axis direction can be easily set to be the same. .

  The sum of the values obtained by dividing the thickness of each propellant by the combustion speed is made constant so that the sum of the combustion times is the same. The nozzle-less rocket according to item. When comprised in this way, the sum of the combustion time of a propellant can be made the same with a simpler parameter | index.

  The nozzleless rocket according to any one of claims 1 to 3, wherein the propellant is a composite propellant. When comprised in this way, the combustion rate of a propellant can be adjusted easily over a wide range.

A sectional view showing a nozzleless rocket of a 1st embodiment which materialized the present invention. (A) is a graph which shows the relationship between a combustion pressure and a combustion rate, (b) is a graph which shows the relationship between the flow velocity of the combustion gas which flows through the surface, and a combustion rate. (A) is a partially enlarged sectional view showing a nozzleless rocket, (b) is a partially enlarged sectional view of a nozzleless rocket when the shape of the second propellant is different, and (c) is a second propellant of (b). The partial expanded sectional view of the nozzleless rocket which shows the middle of combustion in the case of a shape. (A)-(e) is sectional drawing which shows the process of forming a propellant in layers in a combustion case. The graph which shows the relationship between the combustion time and combustion pressure of the nozzleless rocket of 1st Embodiment. Sectional drawing which shows the state which formed the 1st propellant and the 2nd propellant in layers. The graph which shows the relationship between a combustion pressure and a combustion speed about the nozzleless rocket incorporating the 1st propellant and the 2nd propellant of FIG. The graph which shows the relationship between combustion time and a combustion pressure about the nozzleless rocket incorporating the 1st propellant and the 2nd propellant of FIG. Sectional drawing which shows the other example of a nozzleless rocket. Sectional drawing which shows the ducted rocket engine of 2nd Embodiment. The graph which shows the relationship between combustion time and combustion pressure about the ducted rocket engine of FIG. Sectional drawing which shows the conventional nozzleless rocket. FIG. 13 is a graph showing the relationship between the position in the central axis direction and the pressure of the combustion gas flow or the flow velocity of the combustion gas flow for the nozzleless rocket of FIG. 12. (A) is a graph which shows the relationship between the combustion pressure of a propellant, and a combustion rate, (b) is a graph which shows the relationship between the flow velocity of the combustion gas which flows through the surface, and a combustion rate. Explanatory drawing which shows the change of the inner surface shape of the propellant of the nozzleless rocket of FIG. (A) And (b) is a graph which shows the relationship between combustion time and combustion pressure about the nozzleless rocket of FIG. Sectional drawing which shows the ducted rocket engine which has the conventional nozzleless rocket.

Explanation of symbols

  DESCRIPTION OF SYMBOLS 1 ... Combustion case, 1a ... Inner wall surface, 2 ... Ignition device, 3 ... Combustion chamber, 4 ... 1st propellant, 4b ... Inner surface, 5 ... 2nd propellant, 6 ... Center axis, 7 ... Inner hole, 17 ... ducted rocket engine, 18 ... gas generator, 20 ... nozzleless rocket, 23 ... air intake duct.

Claims (3)

An ignition device is provided at the front of the cylindrical combustion case, the combustion chamber in the combustion case is filled with propellant, and an inner hole extending in the central axis direction of the combustion case is formed in the center of the propellant. In the equipped nozzleless rocket,
At least two types of propellant are filled in a layered manner in the radial direction over the entire region or a part of the central axis direction, and the combustion case of the combustion case is formed radially from the inner surface of the propellant that forms an inner hole at each position in the central axis direction. A nozzleless rocket, characterized in that the sum of the combustion times until reaching the inner wall surface is the same. The radial thickness of each propellant comprising at least two propellants is based on the propellant composition, the diameter and length of the combustion chamber filled with the propellant, and the cross-sectional area and length of the bore. The nozzleless rocket according to claim 1, wherein the nozzleless rocket is set. A gas generator for generating fuel gas is provided in the front part, the nozzleless rocket according to claim 1 or 2 is provided in the rear part, and an air intake duct is provided in the side part. A nozzleless rocket characterized by being used as a booster applied to a ducted rocket engine in which fuel gas from a generator and air from an air intake duct are mixed and burned.

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