US3715888A

US3715888A - Hybrid rocket

Info

Publication number: US3715888A
Application number: US00880286A
Authority: US
Inventors: S Massie
Original assignee: Universal Oil Products Co
Current assignee: Honeywell UOP LLC; Universal Oil Products Co
Priority date: 1969-11-26
Filing date: 1969-11-26
Publication date: 1973-02-13
Anticipated expiration: 1990-02-13

Abstract

An improved hybrid rocket, wherein a tube for conducting liquid oxidizer to a solid fuel charge passes through the solid fuel charge and is constructed so as to structurally disintegrate when subjected to the heat produced by the reaction of the solid fuel and liquid oxidizer. Disintegration of the tube proceeds at approximately the same axial rate as does consumption of the solid fuel during operation of the rocket. As a result, the liquid oxidizer is continually supplied to the appropriate location in the solid fuel charge so as to increase the uniformity and thoroughness of burning. The burn time of the rocket may thereby be extended and the thrust may be varied by utilizing a tube of varied cross section and varied wall thickness.

Description

Ut -13 73 UR 397157888 United States Patent [191 [111 3,715,888 Massie i 1 Feb. 13, 1973 HYBRID ROCKET Primary ExaminerSamuel Feinberg [75] Inventor: Stephen Massie, Palatine [IL Att0rneyJ. R. Hoatson, Jr. and Philip T. Liggett [73] Assignee: Universal Oil Products Co., Des [57] ABSTRACT Plaines, lll.

An improved hybrid rocket, wherein a tube for con- Filedl 1969 ducting liquid oxidizer to a solid fuel charge passes {21] AppL No; 880,286 through the solid fuel charge and is constructed so as to structurally disintegrate when subjected to the heat produced by the reaction of the solid fuel and liquid [52] US. Cl ..60/25l, 60/254 oxidizer Disintegration of the tube proceeds at [51] "Fozk 9/06 proximately the same axial rate as does consumption [58] Field Of Search ..60/251, 254 of the Solid fuel during operation of the rocket. As a result, the liquid oxidizer is continually supplied to the [56] References cued appropriate location in the solid fuel charge so as to UNITED STATES PATENTS increase the uniformity and thoroughness of burning. The burn time of the rocket may thereby be extended and the thrust may be varied by utilizing a tube of essums, r. 3,177,657 4/1965 Strauss et al. ..e0/251 vaned cross Seem and vaned wan 3,274,771 9/1966 Lieberman et al. .....60/25l 5 Claims, 5 Drawing Figures 3,097,481 7/1963 Silver ..60/254 PATENTEDFEB 13 I973 N VE/V TOR. Stephen N. Mass/e J9?" ATTO EYS HYBRID ROCKET This invention relates to an improved hybrid rocket. More particularly, the invention involves a tube for conducting liquid oxidizer to a solid fuel charge which passes through the solid fuel charge and is constructed so as to structurally disintegrate when subjected to the heat produced by the reaction of the solid fuel and liquid oxidizer. Disintegration of the tube proceeds at approximately the same axial rate as does consumption of the solid fuel during operation of the rocket. As a result, the liquid oxidizer is continually supplied to the appropriate location in the solid fuel charge so as to increase the uniformity and thoroughness of burning. The burn time of the rocket may thereby be extended and the thrust may be varied by utilizing a tube of varied cross section and varied wall thickness.

A principal disadvantage of hybrid and solid fuel rockets, as compared with liquid fuel rockets, is that rockets operating on solid fuel tend to be somewhat unpredictable and unreliable in operation. This unpredictability is due primarily to the uncontrolled burning in the rocket engine which occurs once the rocket fuel has been ignited. Conventional rockets having a solid fuel utilize either a solid or a liquid oxidizer to sustain combustion of the fuel. A rocket using a solid oxidizer and a solid fuel is a true solid fuel rocket. A rocket utilizing a solid fuel with a liquid oxidizer is commonly termed a hybrid rocket. While hybrid rockets may be controlled in flight more accurately than true solid fuel rockets by regulating the supply of oxidizer to the solid fuel, the degree of control attainable is insufflcient to render this type of rocket at all predictable in operation when compared with a liquid fuel rocket. The reason that solid fuel rockets are unpredictable in operation is that once the solid fuel is ignited, there is no control over irregularities in burning in the rocket except through regulation of the quantity of oxidizer supplied in a hybrid rocket. Erratic and unpredictable thrust occurs because of the different burning patterns that occur in the solid fuel as influenced by a multitude of uncontrollable conditions. For example, intense burning at one point in the fuel charge may structurally isolate an unburned portion of the solid fuel charge to the extent that it becomes dislodged and is carried out of the exhaust of the rocket without having been consumed in combustion. Similarly, uneven burning of the rocket fuel may result in channeling so that liquid oxidizer supplied to the solid fuel is channeled away from unburned portions of the fuel and is discharged from the rocket exhaust without having reacted with the solid fuel. Various attempts have been made to overcome these problems, but none has heretofore proved successful enough to render solid fuel rockets sufficiently predictable in operation.

It is an object of the present invention to increase the thoroughness of burning of the solid fuel in a hybrid fuel rocket. This is accomplished by sustaining combustion at the downstream portion of the solid fuel charge and by preventing premature ignition upstream from the rocket exhaust. Premature burning causes portions of unburned solid fuel to become structurally isolated from the solid fuel charge and from the fuel case, whereupon the unburned solid fuel is dislodged and carried off in the exhaust. This problem commonly exists in solid fuel charges which burn radially outward, that is, outward toward the fuel case and the rocket shell.

Another object of this invention is to increase the degree of control exerted over the burning of the solid fuel in the rocket during flight, and thereby allow the performance of hybrid rockets to be stabilized by proper design. This maintenance of control after ignition is exerted by establishing a high level of uniformity in the burning of the fuel by preventing premature contact of the oxidizer with the solid fuel. The oxidizer is supplied to specific locations within the solid fuel at specific moments during the time of rocket operation, as predetermined during the rocket construction. Erratic burning patterns in portions of the solid fuel charge are thereby prevented from achieving a cumulative effect to cause significant deviations from the predicted flight pattern or other motion pattern in rockets which are not designed for flight. Control of the effects of isolated combustion irregularities promotes the stability and predictability of the operation of the rocket during flight.

In a broad aspect, this invention is, in a hybrid rocket utilizing a solid fuel and a liquid oxidizer for propulsion, the improvement comprising a solid fuel charge encompassing at least one tube means parallel to the rocket axis through which oxidizer is conducted to said solid fuel charge and which is structurally unstable when subjected to the heat produced by the reaction of said solid fuel and said liquid oxidizer, whereby disintegration of said tube means proceeds at approximately the same axial rate as does consumption of said solid fuel during the operation of said rocket. The term axial rate refers to the rates at which the zone of disintegration of the tubes means and the zone of fuel combustion advance from the exhaust section of the rocket toward the rocket nose.

The construction of the rocket of this invention allows the tube means through which oxidizer is supplied to the solid fuel to disintegrate at an axial rate approximately equal to the axial rate of fuel consumption so that the point of oxidizer-fuel contact is at the downstream extremity of the unburned fuel charge. Thisprevents the oxidizer from prematurely reacting with the fuel upstream in the rocket engine, thereby isolating portions of fuel and causing it to be ejected unburned. By the same token, oxidizer is not injected too far downstream from where it would be ejected without having reacted with the fuel.

In one preferred form of the invention, a single tube is aligned with the rocket axis to serve as the tube means. For ease and economy of construction, the solid fuel charge is molded about the tube during the manufacture of the rocket. In one preferred embodiment of this invention, the tube has a star-shaped cross-sectional configuration. While this configuration is normally associated with radial burning solid fuel rockets, it is useful in this invention because radial burning does take place throughout the relatively short distance that exists between the downstream tube end and the downstream extremity of the fuel charge.

In a refinement of this invention, the cross-sectional area occupied by the tube varies, thereby influencing the rate of fuel consumption which, in turn, affects the rocket thrust. Similarly, the thickness of the tube wall may be varied along the tube axis in order to influence the rate of fuel consumption. By varying either the tube cross section or wall thickness, the rocket may be constructed so as to predictably accelerate and decelerate at predetermined intermediate times during the rocket flight. It should be pointed out that at any given location along the tube axis, the tube wall is uniform.

One very important determination which must be made in the construction of the tube of this invention is the determination of compatible substances to be used for the solid fuel, the oxidizer, and the oxidizer supply tube means. While it is important that this determination be made, there is no lack of workable combinations of fuel, oxidizer, and tube structural materials. Some of the conventional liquid oxidizer-solid fuel hybrid propellant systems include: chlorine trifluoridelithium hydride; chlorine trifluoride-lithium; nitrogen tetroxide-beryllium hydride; hydrogen peroxide-beryllium hydride; fluorine-beryllium hydride; oxygen-beryllium hydride; and fluorine-aluminum hydride. Other possible liquid oxidizers in hybrid rockets include: difluoroamine, trifluoroamine, and oxygen difluoride.

Stainless steel, Monel, and nickel are all compatible with those propellant systems utilizing chlorine trifluoride as an oxidizer. The term Monel includes several nickel alloys, all comprising about 66 percent nickel, about 0.12 to 0.18 percent carbon, about 0.60 to 0.90 percent manganese, about 1.00 to about 1.35 percent iron, about 0.15 percent silicon, about 29.5 to 31.5 percent copper, and in some instances small proportions of sulfur, aluminum, or titanium. Pure aluminum is known to be a good container for hydrogen peroxide and dry fluorine may be safely handled in nickel and Monel, since a protective fluoride film quickly develops. Dry nitrogen tetroxide can be used with nickel, aluminum, stainless steel, lnconel, and carbon steel. lnconel includes nickel alloys containing anywhere from about 32 to about 76 percent nickel, about 0.04 percent carbon, about 0.20 to 0.75 percent manganese, from 6.75 to 46 percent iron, from 0.20 to 0.35 percent silicon, from 0.10 to 0.30 percent copper, from to 20.5 percent chromium, and in some instances small quantities of aluminum, titanium, molybdenum, or columbium. Oxygen difluoride can be used with those structural materials compatible with chlorine trifluoride. Difluoroamine and trifluoroamine can be used with borosilicate glass, polyethylene, polypropylene, or those structural materials compatible with chlorine trifluoride.

Specific combinations of oxidizers, solid fuels, and tube structural materials include:

chlorine trifluoride-lithium hydride-Monel tube,

nitrogen tetroxide-beryllium hydride-carbon steel tube, I hydrogen tube, oxygen-beryllium hydride-glass tube, and fluorine-aluminum hydride-nickel tube. These examples are not meant to be inclusive, as other substances may also be used as long as the oxidizer and the solid fuel are suitable as a propellant system for a hybrid rocket, and as long as the tube is constructed of 'a material which is unstable when subjected to the heat produced by the reaction of the liquid oxidizer with the solid fuel. That is, as long as the heat generated by the reaction of the oxidizer and solid fuel is sufficient to melt, disintegrate, or otherwise decompose the structural material of which the tube is constructed. The tube peroxide-beryllium hydride-aluminum should be of a material which disintegrates slightly upstream along the rocket axis from the zone where the most intense combustion occurs. In this manner, oxidizer is supplied to the solid fuel at or just upstream from the fuel currently being consumed in combustion at any given time. The combustion of the solid fuel follows the disintegrating end of the tube as both proceed axially upstream throughout the length of the solid fuel charge. The axial rate of combustion and decomposition must be approximately equal. If the tube disintegrates too readily, that is, if the tube disintegrates well upstream from the zone of combustion at a given point in time, premature contact between the oxidizer and upstream sections of the solid fuel charge will occur. Conversely, if the heat from the reaction of the oxidizer with the solid fuel does not cause the tube to disintegrate until combustion is occurring right at the tube end, the oxidizer will be supplied downstream from the portion of the solid fuel charge which is to be ignited next. This will result in an unnecessary consumption of liquid oxidizer and an incomplete combustion of the solid fuel which will cause a portion of the thrust potential of the rocket to go unrealized.

This invention may be further illustrated inthe accompanying drawings in which:

FIG. 1 is a side view of one embodiment of the improved rocket of this invention partially broken away and in partial section.

FIG. 2 is an end view of FIG. 1 of the rocket exhaust along the lines 2-2.

FIG. 3 is an alternate embodiment of the rocket exhaust of FIG. 2.

FIG. 4 is an alternate embodiment of the rocket exhaust of FIG. 2.

FIG. 5 is an alternative embodiment of a portion of the rocket engine of this invention.

Referring now to FIG. 1 there is shown a simple rocket 1. Rocket 1 is made up of three general sections, namely, a forward section containing a pressurizing gas storage tank 3 at the left end of FIG. 1, a central section containing the liquid oxidizer storage tank 2, and a rocket engine and exhaust section containing the solid fuel charge 11 and the single conducting tube 14 of the tube means of this invention. Tube 14 conducts liquid oxidizer to the solid fuel charge 11 from the oxidizer storage tank 2 while rocket l is in flight.

The tank 3 contains a suitable pressurizing gas, such as helium, under relatively high pressure. This gas is conducted by a conduit 5 through a valve 4 and another conduit 6 to the adjacent end of oxidizer tank 2. When the valve 4 is opened, gas under pressure pressurizes the liquid oxidizer in tank 2. From oxidizer tank 2 there is a central conduit 9 and peripheral conduits 8 which lead to a chamber 7. Within conduits 8 and 9 are rupture disks which retain the oxidizer within tank 2 before launch. Upon pressurization of tank 2, however, the disks in conduits 8 and 9 are ruptured, whereupon liquid oxidizer flows through conduits v8 and 9 into chamber 7 and through control valve 10 toward the tube 14 and the solid fuel combustion charge 11. Since the pressurizing tank 3, the oxidizer tank 2, and the related conduits and valves are conventional, they need not be described in detail herein.

The oxidizer is conducted from the oxidizer tank 2 to the solid fuel via tube 14. Tube 14 is a single tube axially aligned with the rocket 2 and constructed of a material which is structurally unstable when subjected to the heat produced by the reaction of the solid fuel and the liquid oxidizer. When combustion of the solid fuel occurs, that portion of tube 14 in the vicinity of the combustion area disintegrates. For some oxidizer-fuel combinations, ignition occurs spontaneously upon contact of the oxidizer with the fuel. Other oxidizer fuel combinations require a conventional ignition means to initiate combustion. Such an ignition means is not illustrated in the drawings. Disintegration of the tube 14 proceeds at approximately the same axial rate as does consumption of the solid fuel charge 11 during the operation of the rocket. Tube 14 disintegrates slightly upstream (toward the left along the axis of rocket 1) from the point of most intense combustion at any given time. The disintegration of tube 14 thereby precedes the most intense zone of combustion in the solid fuel charge 11 axially upstream at a relatively short and a relatively fixed'axial distance. Since the oxidizer is able to contact the solid fuel charge 11 in that area in which the tube 14 has disintegrated, the liquid oxidizer is supplied to the solid fuel charge 11 just upstream from where existing combustion is taking place. This results in a steady uniform fuel consumption from the rear of the solid fuel charge 11 in exhaust section towards the front of the solid fuel charge 11 adjacent the valve 10. The oxidizer from tank 2 is prevented from prematurely contacting the solid fuel charge 11 because of the separation provided between the oxidizer and the solid fuel by the wall of tube 14.

As is conventional, the solid fuel charge 11 is molded within a resin impregnated fiberglass wound fuel case 12. Prior to molding, however, the tube 14 is axially positioned within the engine case 12. The solid fuel is thereby molded uniformly about the tube 14 within the rocket engine case 12.

Referring now to FIG. 3, there is shown an alternative embodiment of the tube means of this invention. In FIG. 3, the tube means is a single tube 14 constructed with a star-shaped cross-sectional configuration. This configuration assists in maintaining a uniform radial burn rate once the solid fuel has been ignited downstream from the end of the disintegrating tube means 14. The solid fuel charge 11' of FIG. 3 is molded within rocket engine case 12 and is positioned within the rocket shell 13 in exactly the same manner as is the solid fuel charge 11 of FIG. 2. The same is true of the solid fuel charge 11" of FIG. 4. In FIG. 4, however, the tube means of this invention is comprised of three individual tubes 16. Tubes 16 are parallel to the rocket axis, but none of the tubes 16 is axially aligned with the rocket axis. That is, none of the tubes has an axis which is aligned with or coincides with the rocket axis. The various cross-sectional configurations of the solid fuel charges in FIGS. 2, 3, and 4, are given for purposes of illustrating the conventional configurations, and are not intended to be restrictive as to the scope of this invention.

FIG. 5 illustrates a specialized modification of the tube means and solid fuel charge of this invention. The tube means of FIG. 5 is unique in that the cross-sectional area occupied by the tube 18 varies along the tube axis as does the thickness of the tube wall. These variations influence the rate of fuel consumption in the rocket and make possible predictable periods of acceleration and deceleration before the total consumption of the fuel charge 17. The extreme downstream portion 19 of the tube 18 is constructed with a very thin tube wall. Upon the initial ignition of the fuel charge 17, this extreme downstream portion 19 of the tube 18 will rapidly disintegrate, thereby exposing a substantial area of the fuel charge 17 to the oxidizer coming from the upstream portion of the tube 18. After the extreme downstream portion 19 of tube 18 has disintegrated, the rate of fuel combustion decreases somewhat because of the thickened walls of the portion 20 of the tube 18. The thicker walls of portion 20 require greater heat for total disintegration, so that the distance along the rocket axis between the downstream edge of the tube 18 and the zone of most intense combustion of the fuel charge 17 is much shorter than it was after initial fuel ignition. The result is that less fuel is exposed to the liquid oxidizer than was the case when the rocket was first ignited so that the combustion rate and the rocket thrust are decreased. As the combustion proceeds upstream, the thick-walled portion 20 gives way to a portion 21 of the tube 18 having still thicker walls. As the area of most intense combustion in fuel charge 17 reaches the vicinity of the portion 21 of tube 18, the increased thickness of the tube walls at portion 21 slows disintegration of the tube in this area until the zone of most intense combustion of the fuel charge 17 is practically in the same general plane perpendicular to the rocket axis as is the downstream extremity of the tube 18. During this portion of the rocket flight, the fuel consumption is minimal and the rocket will decelerate due to the forces of gravity and air friction. Once past this thickened portion 21, however, the thickness of the tube walls abruptly decreases and the cross-sectional area of the tube 18 widens at the portion 22 of tube 18. The increased cross-sectional area and the thin tube walls causes the rate of consumption of the fuel charge 17 to markedly increase, since there is greater timely contact of the oxidizer with the unburned fuel. During the time the fuel charge 17 is being consumed in the vicinity of portion 22 of tube means 18, the rocket will accelerate due to the increased fuel consumption. At portion 23 of tube 18, the tube walls again thicken and the cross-sectional area occupied by the tube decreases, thereby decreasing the rate of fuel consumption. This transition is quite similar to the transition that occurred when the portion 19 of tube 18 had disintegrated and the slower more time consuming disintegration of portion 20 of tube means 18 was begun.

It is to be understood that the foregoing descriptions of several of the embodiments of this invention are intended to be illustrative only, and that numerous conventional alterations in the rocket design can be made without departing from the spirit of the invention as defined in the claims.

I claim as my invention:

1. A hybrid rocket utilizing a solid fuel and a liquid oxidizer for propulsion comprising a solid fuel charge encompassing at least one tube means, the thickness of said tube means wall varying along the tube means axis, said tube means being disposed parallel to the rocket axis through which tube means oxidizer is conducted to said solid fuel charge and which tube means is structurally unstable when subjected to the heat produced by the reaction of said solid fuel and said liquid oxidizer, whereby disintegration of said tube means proceeds at approximately the same axial rate as does consumption of said solid fuel during the operation of said rocket.

2. The rocket of claim 1 further characterized in that said liquid oxidizer is chlorine trifluoride, said solid fuel is lithium hydride, and the aforesaid tube means is constructed of Monel.

3. The rocket of claim 1 further characterized in that said liquid oxidizer is hydrogen peroxide, said solid fuel is beryllium hydride, and the aforesaid tube means is constructed of glass.

4. The rocket of claim 1 further characterized in that said liquid oxidizer is fluorine, said solid fuel is aluminum hydride, and the aforesaid tube means is constructed of nickel.

5. A hybrid rocket utilizing a solid fuel and a liquid oxidizer for propulsion comprising a solid fuel charge encompassing at least one tube means parallel to the rocket axis, the major portion of said tube means having a star-shaped cross-sectional configuration, through which tube means oxidizer is conducted to said fuel charge and which tube means is structurally unstable when subjected to the heat produced by the reaction of said solid fuel and said liquid oxidizer, whereby disintegration of said tube means proceeds at approximately the same axial rate as does consumption of said solid fuel during the operation of said rocket.