US3650214A

US3650214A - Toy rocket and gas propellant system

Info

Publication number: US3650214A
Application number: US814484*A
Authority: US
Inventors: Charles J Green
Original assignee: VASHON IND Inc
Current assignee: VASHON IND Inc
Priority date: 1969-01-21
Filing date: 1969-01-21
Publication date: 1972-03-21
Anticipated expiration: 1989-03-21

Abstract

A model rocket propelled by a non-flammable liquefied gas is disclosed which is capable of realistic performance as well as appearance. The rocket has (1) a novel structural arrangement for separation of the nose portion of the rocket from the rocket body once the gas pressure in the pressure chamber has dropped and thrust is dissipated, (2) a novel valving arrangement for initiating the thrust which may be triggered remotely by means of an electrical current, and (3) a liquefied gas propellant which forms a gas-liquid mixture as it flows through the rocket nozzle, the specific impulse of the gas-liquid mixture being considerably higher than the specific impulse obtainable from the pure liquid.

Description

D United States Patent [151 3,650,214

Green 1 Mar. 21, 1972 s41 TOY ROCKET AND GAS PROPELLANT 2,918,751 12/1959 Johnson ..102/34 x SYSTEM 3,029,704 4/1962 Truax ..l02/49.3 3,082,666 3/l963 Fitzpatrick. ....60/254 [721 vashm Island wash $266,238 8/1966 Wilkerson ..102/34 x Assignee; vashon Indus'ries, Inc vashon Island 3,3 l 3,1 MBXSOXI et al X Wash. Primary ExaminerVerlin R. Pendegrass [22] 1969 Attorney-Christensen & Sanborn 211 Appl. No.: 814,484

[57] ABSTRACT Related US. Application Data A model rocket r0 elled by a non-flammable li uefied as is P P q 8 commuanon'm'pan of 700,822, 26, disclosed which is capable of realistic performance as well as 1968' abandonedappearance. The rocket has (1) a novel structural arrangement for separation of the nose portion of the rocket from the t 4 s s i i i i s f dropped and thrust is dissipated, (2) a novel valving arrange- [58] Fleld of Search ..l02/34, 34.1, 493604354; mm for initiating the thrust which may be triggered remotely by means of an electrical current, and (3) a liquefied gas propellant which forms a gas-liquid mixture as it flows through [56] References Cited the rocket nozzle, the specific impulse of the gas-liquid mix- UNITED STATES PATENTS ture 1611}; conrsliderablly higiher than the specific impulse 0btama e rornt e ure 1 U1 2,779,283 1/1957 Baughman ..l02/49.4 p q 2,787,218 4/1957 Anthony 102/494 11 Claims, 11 Drawing Figures 62% 63 1 l 7 6i f4 21; 167 I 5e\ w I fie \7 66 u l3 le' l 24 I, 23 2| I9 22 l.

w 27 I i- N l 26 V l 34 l l 36 l 37 PATENTEUMARZI x972 SHEET 1 [IF 3 FlG 2 V FIG. I

INVENTOR. CHARLES J. GREEN ATTORNEYS PAIENIEnmz: m2 3. 650,2 1 4

sum

3 or 3 INVENTOR. GHARLES J. GREEN ATTORNEYS TOY ROCKET AND GAS PROPELLANT SYSTEM BACKGROUND OF THE INVENTION The present application is a continuation-in-part of my application Ser. No. 700,822, filed Jan. 26, 1968, and titled Toy Rocket. The said earlier filed application now having been abandoned in favor of the present application.

This invention relates to rockets and reaction propulsion devices, particularly adapted for use as toys or models wherein the thrust is obtained by the use of a liquefied gas or solutions or emulsions of liquefied gases contained within a pressure vessel. The fluid within the pressure vessel is allowed to flow through a noule with the driving pressure being provided by the fluids own vapor pressure.

Although various designs of aerial toy devices and rockets are available, most such devices are inadequate, ineffective, or considered to be unsafe because of fire and/or explosion hazard. The present invention encompasses a rocket device preferably utilizing a non-flammable, nontoxic liquefied gas which provides a safe, low-cost and effective rocket device suitable for use by children, modelers and hobbyists. The rocket device of the present invention is adapted for both vertical and horizontal staging, may be actuated from a remote position and provides a high degree of safety, realism, and flexibility. The invention also encompasses a novel structure arrangement for separation of the nose cone portion of the rocket from the rocket body once the gas pressure has dropped and the thrust is dissipated. Likewise in vertical staging of a plurality of propellant charges the gas pressure is utilized to maintain the sections connected and separation occurs upon destruction of the gas pressure. A novel valving arrangement is also provided for initiating the thrust and may be triggered remotely by means of an electrical current or may be done manually if desired.

Although the present invention is described in terms of a model toy device, it will be understood that the structure and concept may be utilized for rocket devices of a greater size and range for other purposes.

Accordingly, the object of the present invention is to provide an improved, low-cost, safe and effective rocket device utilizing a relatively low-pressure liquefied gas as a propellant charge.

Another object of the present invention is to provide a novel liquefied gas propellant which is in gas form at ambient temperature and atmospheric pressure, the propellant, on exhaust, forming an gas-liquid mixture in the rocket nozzle.

Another object of the present invention is to provide a rocket device of the character described which is readily adapted for staging, both vertically and horizontally.

Another object of the present invention is to provide a rocket device of the character described wherein the propellant gas pressure is used to control the separation of the nose cone section as well as tandem rocket stages.

Another object of the present invention is to provide a rocket device of the character described which is simplified in structure and easily and safely chargeable with a propellant charge so as to be suitable for use by children, modelers and hobbyists.

A further object of the present invention is to provide a rocket device ofthe character described wherein the liquefied gas is allowed to flow through a nozzle and wherein a novel valving arrangement, adapted for remote electrical triggering, is used for the nozzle.

A still further object of the present invention is to provide a rocket device of the character described wherein the novel valving arrangement for initiating thrust may also be utilized for successive firing of horizontally staged or clustered" rocket units.

Other objects and advantages of the present invention will be apparent to those skilled in the art from the following specification and claims and from the accompanying drawings wherein:

DESCRIPTION OF THE DRAWINGS FIG. 1 is a side-elevational view of a single stage rocket device in firing position according to the present invention;

FIG. 2 is a cross sectional view taken along lines 22 of FIG. 1;

FIG. 3 is a top plan view of the rocket device in firing position;

FIG. 4 is a partially sectioned elevational view of the nose cone illustrating the position of a parachute device;

FIG. 5 is a partially sectioned elevational view of a propellant storage container;

FIG. 6 is a sectional view, taken along line 6-6 of FIG. 7, of the separator section utilized between rocket sections during vertical staging;

FIG. 7 is a cross sectional view taken along lines 7-7 of FIG. 6 and showing the separator section in place between two tandem rocket sections;

FIG. 8 is a side elevational view of a cluster of rocket devices in the firing position and arranged for horizontal stagmg.

FIG. 9 is a cross sectional view taken along lines 9-9 of FIG. 8;

FIG. 10 is a cross sectional view of a switch mechanism in the area of the circle in FIG. 8; and

FIG. 11 is cross sectional view of the preferred structural arrangement for separation of the nose cone from the rocket body.

DESCRIPTION OF THE INVENTION The basic single stage rocket device will first be described relative to FIGS. 1 through 5 of the drawings. The rocket device comprises a rocket body indicated generally at l which is a hollow cylindrical body comprising the rocket motor or propellant charge container. The rocket body is provided with conventional fins 2 of which there are three shown in FIGS. 1 and 3 and a nose cone section 3. A separator assembly 4 serves to connect the nose cone section and the rocket body during flight and until the propellant charge is exhausted. In order to provide thrust, the liquefied gas propellant is allowed to flow through a nozzle indicated generally at 6 at the bottom end of the rocket body 1 which has a releasable nozzle plug indicated generally at 7 for initiating thrust. When readied for firing, the rocket device is positioned on a vertical guide rail 8 mounted on the base 9 as shown in FIG. 1.

The body portion 1 is an elongated cylindrical tubular container which can be constructed from any lightweight lightgauge metal of sufficient strength to withstand internal gas pressures as required. The upper end of the body 1 is provided with a dished closure 11 with the outer peripheral edge thereof seating against a shoulder 12 formed by the inwardly directed lip portion of the body. An adhesive 13 of a suitable type may be used to seal the end closure 11 in position as illustrated in FIG. 2. An opening in the center of the closure 11 has an externally threaded fitting 14 which has a through passage and a tapered seat 16 for receiving an O-ring seal 17 for a purpose to be described. A conventional pressure release valve assembly indicated generally at 18, is located in the rocket body and serves to prevent the gas chamber from being overpressurized when filling. The pressure release valve assembly 18 includes a valve body 19 which extends through the wall of the rocket body 1 and includes a head which seats against the wall of the rocket body as illustrated in FIG. 2. The valve stem 19 is surrounded by a spring 21 which seats at one end against the stop collar 22 and against the other end on a metal washer 23 with a rubber sealing washer 24 providing a seal between the wall of the rocket body and the head of the valve stem 19 until internal pressure acting on the valve stem compresses the spring 21 to unseat the valve. For filling purposes a standard filling valve insert indicated generally at 26 is located in the wall of the rocket and comprises a Neoprene body 27 held in place by the clips 28 within the rocket body. The assembly 26 is a standard item available from a number of sources. The filling valve 27 is designed to receive a conventional filling needle 29, shown in FIG. 5, for dispensing the liquefied gas under pressure from a pressurized container 31 which is also of conventional design and commercially available for dispensing pressurized gases and liquids.

The bottom end of the rocket body 1 has a closure 32 which may be identical to the closure 11 and may be installed in the same manner. The central opening in the closure 32 is fitted with the nozzle member 6 as illustrated. The nozzle structure 6 has a through passage 33 with a diverging outlet. The lower end of the nozzle member 6 is provided with an external conical taper 34 which matches the taper 36 on the sleeve or nozzle extension 37. Both the nozzle structure 6 and the extension 37 may be made from metallic material with the

tapers

34 and 36 being designed so as to provide a frictional holding engagement between the nozzle and the extension. The metallic nozzle plug 7 is inserted within the passage 33 and includes a conventional O-ring seal 38 for sealing the gas pressure chamber provided by the rocket body 1. The plug 7 is normally held in place by the triggering mechanism presently to be described and is allowed to be blown out by the internal gas pressure to fire the rocket and to initiate thrust. As a safety feature a removable lock pin 37a may be provided in the sleeve 37.

The thrust initiator or firing device involves the use of a bimetallic structural element 39 which, in the present embodiment, is in the form of a wire, but which could conceivably be in the form of foil or other shape. The bimetallic element 39 and the frictional engagement between the nozzle 6 and sleeve 37 restrain the plug 7 from being blown out by the internal gas pressure within the rocket body 1. The bimetallic element consists of aluminum and platinum or suitable alternative material, which when heated above a critical temperature, form an alloy accompanied by a rapid release of heat. The alloying action and the accompanying heat release causes the bimetallic element to lose its structural integrity and the valve plug 7 is permitted to be blown out by the gas pressure. The bimetallic element is available commercially and known to the art, one trade name of such a typical material is Pyro-Fuse. In the present embodiment, the bimetallic element 39 is placed between the

tapers

34 and 36 when the sleeve 37 is applied to the nozzle 6. The lower end of the plug 7 likewise has a taper 41 which is designed to cooperate with the internal taper 42 of an electrical cap 43. The cap 43 is connected to the electrical lead wire 44 by means of the terminal member 46. A second lead wire 47 has the electrical conductor thereof inserted within the bottom end of the plug 7 as illustrated. The taper 41 on the lower end of the plug 7 is provided with a covering of dielectric material 48 which electrically insulates the nozzle plug from the conductor cap 43. The bimetallic member 39 extends downwardly between the inside tape 42 of the cap 43 and the dielectric sleeve and may be simply wound about the insulation on the lead wire 47 as illustrated. When an electric current is passed through the lead wire 44 the current is caused to flow through the bimetallic member 39, then via the nozzle 6 to the plug 7 and to the lead wire 47 to complete a circuit. The resulting destruction of the bimetallic restraining element allows the gas pressure in the rocket body to blow the plug 7 from the nozzle and thrust is initiated.

The separator assembly 4 serves to connect the nose cone section 3 to the rocket body 1 during flight and to cause separation of the two sections after the gas pressure in the body 1 has sufficiently decayed and vertical thrust has been exhausted. FIGS. 2 and 11 illustrate the details of the separator assembly which will also be in the form ofa hollow tubular section of the seam outside diameter as the rocket body 1. The assembly 3 is provided with oppositely facing dished

end closures

49 and 51 which are fitted to the separator section in the same manner described relative to the

end closures

11 and 32 of the rocket body. The lower closure 49 has an externally threaded fitting 52 which is identical with the fitting 14 with the ring 53 serving to connect the separator section with the rocket body as shown. As shown in FIG. 2 the fitting 52 also has a tapered seat 54 for contacting the O-ring seal 17 when the two parts are tightened down. The upper closure member 51 has a central opening to allow the button 56 to reciprocate therethrough. The button 56 is seated on a flexible diaphragm 57 which is positioned against a metal plate 58. The plate 58 also has a central opening 59 to allow gas pressure to act upon the diaphragm in a manner presently to be described. The nose cone section 3 is also a hollow cylindrical structure having the same outside diameter as the rest of the rocket. The nose cone is held to the separator section 4 by means of the leaf spring gripping element 61 which is actuated by the movable button 56. The opposite ends of the spring 61 are in contact with the underside of a stationary plate 62 connected to the separator structure by means of the clip 63. The ends of the spring 61 pass through suitable slots or openings in the clip 63 in such a manner that, as the button 56 is moved upwardly, from the position shown in FIG. 2, the spring 61 will flex and the ends will be pushed outwardly so as to contact the inside surface of the nose cone section 3. In this manner, as long as the button 56 is actuated by upward flexing of the diaphragm 57, the nose cone will be connected to the separator section 4. When the diaphragm is allowed to return to the condition shown in FIG. 2, the leaf spring gripping member 61 is retracted allowing the nose cone to separate. The hollow nose section 3 may be used to store a suitable parachute 64 which is connected to the nose cone as well as to the plate 62 carried by separator section as shown in FIG. 4.

In order to assure that the nose cone section will not separate until a given time lapse, after the loss of the gas pressure within the rocket body 1, a restricted passage for slow bleeding of pressure from the separator section to the chamber 1 and vice versa is provided. The restricted passage may be in the form of a hollow needle 66 placed inside O-ring 17 as shown in FIG. 2, or in the form of a paper disc or discs 66a permeable to the propellant used, as shown in FIG. 11. The needle 66 may be provided with a removable pin 67 to further retard the pressure bleed.

As shown in FIGS. 1 and 3, the rocket body is provided with the

guides

68 and 69 which engage the vertical rail 8 and which serve to initially direct the rocket vertically. The guide 69 may rest on a suitable stop 71 on the rail 8 to support the rocket in firing position.

Rocket performance is principally determined by two characteristics: (1) specific impulse (largely a function of how energetic the rocket propellant is) and (2) vehicle mass ratio (ratio of total loaded weight to empty or burnout weight). Although in large rockets it is desirable to maximize the values of the above mentioned parameters, this in not easily done, nor is it necessarily desirable with model or toy rockets. Rocket propellants having a high specific impulse are generally hazardous to use, particularly for children. Specific impulse is a term used to designate the pounds of thrust which can be generated by a rocket propellant divided by the number of pounds per second of propellant consumed in generating the thrust. The specific impulse of rockets used by the military and NASA is generally in excess of 200 seconds. The specific impulse of the preferred propellant used in the rocket of this invention is about 9 seconds. While this is relatively low compared with the specific impulse of the propellant used in military and NASA rockets, it is desirable in the model rocket of this invention to achieve realistic performance, i.e., low acceleration at lift off and high acceleration at burnout.

To compensate for the relatively low specific impulse of the liquefied gas propellant, the rocket of this invention is structured to have a high mass ratio so that it is capable of reaching high altitudes and performing in a manner similar to rocket vehicles launched by NASA. Theoretically when the mass ratio of the rocket has a value up to 2.718, the vehicle is capable of traveling as fast as its own exhaust gases, (under drag free 0 g. conditions). If the mass ratio is greater than 2.718, vehicle velocity can be greater than the exhaust gas velocity. The rocket of this invention can achieve mass ratios greater than 2.718 with the preferred propellant; however, its maximum velocity is considerably lower than the velocity of high mass fraction military or NASA rockets due to the lower specific impulse of the propellant. The preferred propellant used in the rocket of this invention is a fiourinated hydrocarbon which has a relatively low specific impulse, is nontoxic, noncombustible, has a vapor pressure of approximately 100 p.s.i. at room temperature, has a relatively low molecular weight, and can be stored in low-cost refillable pressure containers. Specifically the preferred fiuorinated hydrocarbon is diehlorodifiuoroethane (Freon-l2, a product of IE. DuPont deNemours & Company). Other fluorinated hydrocarbons such as Freon-22 may be used.

By using liquefied gases which have such reasonably low vapor pressures, light-weight, low-cost pressure vessels can be employed, additionally, low vapor pressures facilitate safe refilling of the pressure vessel by untrained users. Other material may be used in conjunction with the liquefied gas for thermodynamic reasons (improvement in performance) or to provide a more spectacular visible display i.e., smoke generation).

The rocket is made ready for firing by inserting the plug 7 in position so as to be held by the bimetallic elements 39 as described. The safety pin 37a may be used until the rocket is ready for firing. Either the hollow needle 66 with its O-ring seal 17 or the paper discs illustrated in FIG. 11 is put in place and the fitting 52 is tightened into ring 53 to compress the O- ring or paper disc. The nose cone is then set in place on the separator section 4 as illustrated in FIG, 2. The propellant chamber of the rocket body 1 is then filled by inserting the needle 29 into the filling valve 27 in a conventional manner. The rocket body may only be pressurized to a predetermined point at which the release valve 19 will open. The pressure in the body 1 then slowly bleeds through the needle 66 or the paper discs 66a and acts on the diaphragm 57 to operate the button 56 and expand the gripping spring 61 thus connecting the nose cone to the rest of the rocket. When the rocket is in position as shown in FIG. 1, the pin 37a is removed. An electrical current is passed through the

lead wires

44 and 47 to destroy the bimetallic element 39, thereby releasing the plug 7 and initiating thrust. The propellent is a liquid as long as it is confined in the propellant chamber under its own vapor pressure. When plug 7 is released the liquid propellant is forced through the nozzle throat 33 of the nozzle 6 by its own vapor pressure. In the course of its flow through the nozzle the free stream fluid pressure is reduced, causing a portion of the liquefied gas to vaporize. The resulting vapor and liquid mixture is exhausted from the diverging nozzle at high velocity, providing rocket thrust. The specific impulse of the gas-liquid mixture is considerably higher than the specific impulse attainable from the pure liquefied gas and somewhat lower than that attainable from the pure vapor. Once outside the nozzle, the surrounding air rapidly vaporizes the remainder of the liquid propellant.

After the rocket has ascended and pressure within the rocket body 1 drops, the gas pressure in the separator assembly will bleed back through the needle 66 or the paper discs 66a, allowing the diaphragm 67 to return to its relaxed position, withdrawing the gripping spring 61, at which time the nose cone will fall away from the separator assembly. The parachute 64 will be withdrawn from the nose cone due to differences in aerodynamic drag between the separated sections of the rocket.

FIGS. 6 and 7 illustrate the arrangement whereby successive propellant chambers or stages may be vertically arranged in tandem to boost the range of the rocket. As shown in FIG. 7, the lower rocket body 71 is identical in all respects to the rocket body I for the single stage except that the stationary plate 72, which corresponds to the stationary plate 62 shown in FIG. 2, is provided with a nozzle plug 73 which is similar to the plug 7. The second stage or propellant chamber 74 may be identical to the previously described rocket body 1 and is provided with the nozzle structure 76 fitted to the lower end closure in the manner previously described. A separator assembly 77 which is a hollow cylindrical section having the same diameter as the

sections

71 and 74 has its upper peripheral edge provided with bent-in tabs 78 which cooperate with the nozzle extension 79 to hold the separator section 77 to the upper rocket stage 74 when the extension 79 is frictionally engaged to the nozzle 76. In this manner, the separator section 71 becomes an extension of the rocket body 74. The lower section 71 is connected to the separator section 77 by means of the leaf spring gripper element 81 which is subject to the gas pressure within the lower section in the same manner as described with relation to the gripper element 61 in FIG. 2. As the gas pressure in the successive tandem propellant chambers is exhausted, the associated gripping element will be released allowing the next stage to blow its nozzle plug 72 to separate the spent stage. As many stages as practical may be connected in tandem in this manner.

FIGS. 8 to 1 0 illustrate an arrangement whereby a plurality '7 I of rocket devices may be clustered" or horizontally staged. In this arrangement, the two

sections

82 and 83 on the same horizontal level may be identical in structure to the rocket device shown in FIGS. 1 through 4 with the addition of a switch operator 84 as shown in FIG. 10 being connected to the reciprocating button 56. The switch operator 84 serves to break contact between the leaf spring switch arm 86 and the terminal 87. As shown in FIG. 9, the nose cone of one or both of the

rockets

82 and 83 will be provided with

dry cell batteries

88 and 89 connected to the

lead wires

91 and 92 which are used to fire the third or center rocket 93. As will be noted, when gas pressure operates on the button 56, the switch operator 84 holds the switch arm 86 open and, when the pressure is lost in the two

outside rockets

82 and 83, the switch arm will be allowed to close so as to fire the center rocket 93 in the same manner as described relative to FIG. 2. The center rocket 93 is provided with upper and lower sets of

guide sleeves

94 and 96 respectively which align with upper and lower guide sleeves 97 and 98 on the two outside rockets. Slip pins 99 are inserted within the aligned sleeves to hold the rockets in position. The two outside rockets will initially be supported on their extended fins as illustrated and a guide rail 101 may be used to initially direct the cluster vertically. In flight, when thrust is lost by the two outside rockets and the inner rocket is fired, the rockets will separate allowing the two outside rockets to fall away and the inner rocket 93 to contmue.

From the foregoing it will be obvious to those skilled in the art that the present invention provides significant improvements in toy rocket devices of the character described in the nature of safety and efficiency as well as providing a high degree of realism to the user. The combination of low energy or low specific impulse propellant and the true-to-scale high mass fraction enables the rocket of this invention to reach altitudes of up to 1,000 feet. The rockets can also be staged to carry larger payloads or to achieve greater altitudes. The arrangement and types of structural components utilized within this invention may be subjected to numerous modifications well within the purview of the invention and applicant intends only to be limited to a liberal interpretation of the specification and appended claims.

The embodiments of the invention in which an exclusive property or privilege is claimed are defined as follows:

1. An aerial rocket device having a tubular rocket body with a propellant pressure chamber for containing a liquefied gas propellant, said propellant chamber having a nozzle in one end thereof providing a passage for expanding pressure fluid from the propellant chamber to provide thrust, a nozzle plug, means for holding said plug in closing position in said nozzle, and selectively operable release means for releasing said plug to initiate thrust, said propellant chamber being provided with filling valve means for receiving liquefied gas propellant charge from a pressurized storage container, said release means for holding said plug in closing position comprising a destructible restraining element connected between said plug and the nozzle with sufficient strength to withstand the gas pressure within the propellant chamber tending to blow the plug from the nozzle, said nozzle extending beyond the wall of said propellant chamber, a tubular sleeve member adapted to be engaged on the extended portion of the nozzle for holding said destructible restraining member to the nozzle, said plug having a portion thereof extending beyond said nozzle, and a cap member adapted to engage the extended portion of the plug for holding said destructible restraining member to the plug.

2. The rocket device according to claim 1 wherein; said destructible restraining member comprises a bimetallic element, the metals of which alloy upon the passage of electrical current with a resultant loss of structural integrity whereby the plug is released from the nozzle, and means to pass an electrical current through the bimetallic element.

3. The rocket device according to claim 1 including; a nose cone section and a separator section, said propellant chamber and said nose cone and separator sections being matching tubular cylindrical members, second pressure chamber means in said separator section, means to connect said separator section with the propellant chamber with the second pressure chamber being subjected to the gas pressure in the propellant chamber, and gripping means responsive to the gas pressure in the second pressure chamber to normally hold the nose cone section to the separator section until the pressure drops below a predetermined value.

4. The rocket device according to claim 3 wherein said second pressure chamber is provided with a flexible wall for transmitting pressure, said gripping means comprising; a stationary bearing member connected to the separator section and extending into the nose cone, a movable spring having terminal ends adapted to contact the bearing member and to be guided radially thereby to engage the surface of the nose cone section upon flexing of the spring, and actuator means in contact with said flexible wall and said spring for transmitting motion of the flexible wall to the spring, whereby said spring is flexed for engaging the nose cone when the flexible wall is acted upon by the gas pressure within said second pressure chamber.

5. The rocket device according to claim 4 including; means providing a restricted orifice between said propellant chamber and said second pressure chamber for retarding the pressure drop in the second pressure chamber upon loss of gas pressure in the propellant chamber.

6. The rocket device according to claim 5 including a disc or discs between said propellant chamber and said second pressure chamber for retarding the pressure drop in the second pressure chamber upon loss of gas pressure in the propellant chamber, the disc or discs being permeable to the flow of propellant gas therethrough.

7. An aerial rocket comprising a plurality of devices connected in tandem and having leading and trailing ends to form separable rocket sections with one of the sections comprising the trailing section, each device having a tubular rocket body with a propellant pressure chamber for containing a liquefied gas pro ellant, said propellant chamber having a nozzle in one end thereof providing a passage for expanding pressure fluid from the propellant chamber to provide thrust, the leading end of each device having a nozzle plug fixed thereto and engaged in the nozzle of the trailing end of the next successive section; each said section having gas pressure responsive holding means on its leading end for holding the sections together until the gas pressure in the propellant chamber thereof drops below a predetermined level; said last section including a nozzle plug and means for holding said plug in closing position in said nozzle, and selectively operable release means for releasing said plug of said last section to initiate thrust; said propellant chambers being provided with filling valve means for receiving a liquefied gas propellant charge from a pressurized storage container; said release means for holding said plug in closing position comprising a destructible restraining element connected between the associated plug and the nozzle with sufficient strength to withstand the gas pressure within the propellant chamber tending to blow the plug from the nozzle.

8. The combination according to claim 2 wherein a plurality of said devices are connected in tandem and having a leading and trailing end to form separable rocket sections with one of the sections comprising the trailing section, the leading end of each device having a nozzle plug fixed thereto and engaged in the nozzle of the trailing end of the next successive section; each said section having gas pressure responsive holding means on its leading end for holding the sections together until the gas pressure in the propellant chamber thereof drops below a predetermined level, said selectively operable release means being operatively associated with the plug and nozzle of the trailing section.

9. The combination according to claim 7 wherein said gas pressure responsive holding means comprises; a second pressure chamber in pressure communication with the propellant chamber and having a flexible wall for transmitting pressure, a stationary bearing member connected to the rocket section, a movable spring having terminal ends adapted to contact the bearing member and to be guided radially thereby to engage a portion of the trailing end of the next successive rocket section upon flexing of the spring, and actuator means in contact with said flexible wall and said spring for transmitting motion of the flexible wall to the spring, whereby said spring is flexed for engaging the next successive rocket section when the flexible wall is acted upon by the gas pressure within said second pressure chamber.

10. The combination according to claim 2 wherein a plurality of said devices are arranged in a cluster with some of the devices providing an initial thrust stage and the remaining devices providing subsequent thrust stages, means for supporting the subsequent thrust stage devices on the initial thrust stage devices, said supporting means serving to maintain the devices in fixed relation in vertical flight as long as thrust is obtained from the initial thrust stage devices, a source of electrical current carried by at least on of said initial thrust stage devices, electrical conductor means for completing a circuit through the destructible restraining element of the subsequent thrust stage device and said source of electrical current, switch means in said circuit, means responsive to the gas pressure in said one initial thrust stage device for normally holding said switch open until said gas pressure drops below a predetermined level.

11. The combination according to claim 3 wherein a plurality of said devices are arranged in a cluster with some of the devices providing an initial thrust stage and the remaining devices providing subsequent thrust stages, means for supporting the subsequent thrust stage devices on the initial thrust stage devices, said supporting means serving to maintain the devices in fixed relation in vertical flight as long as thrust is obtained from the initial thrust stage devices, a source of electrical current carried by at least one of said initial thrust stage devices, electrical conductor means for completing a circuit through the destructible restraining element of the subsequent thrust stage device and said source of electrical current, switch means in said circuit, means responsive to the gas pressure in said one initial thrust stage device for normally holding said switch open until said gas pressure drops below a predetermined level.

Claims

7. An aerial rocket comprising a plurality of devices connected in tandem and having leading and trailing ends to form separable rocket sections with one of the sections comprising the trailing section, each device having a tubular rocket body with a propellant pressure chamber for containing a liquefied gas propellant, said propellant chamber having a nozzle in one end thereof providing a passage for expanding pressure fluid from the propellant chamber to provide thrust, the leading end of each device having a nozzle plug fixed thereto and engaged in the nozzle of the trailing end of the next successive section; each said section having gas pressure responsive holding means on its leading end for holding the sections together until the gas pressure in the propellant chamber thereof drops below a predetermined level; said last section including a nozzle plug and means for holding said plug in closing position in said nozzle, and selectively operable release means for releasing said plug of said last section to initiate thrust; said propellant chambers being provided with filling valve means for receiving a liquefied gas propellant charge from a pressurized storage container; said release means for holding said plug in closing position comprising a destructible restraining element connected between the associated plug and the nozzle with sufficient strength to withstand the gas pressure within the propellant chamber tending to blow the plug from the nozzle.