EP0269642A1

EP0269642A1 - Rocket propellant charge with negative pressure exponent

Info

Publication number: EP0269642A1
Application number: EP87901295A
Authority: EP
Inventors: Robert David Irvine; Peter Clive Winch
Original assignee: Commonwealth of Australia; Commonwealth of Australia Department of Defence
Current assignee: Commonwealth of Australia; Commonwealth of Australia Department of Defence
Priority date: 1986-02-28
Filing date: 1987-02-26
Publication date: 1988-06-08
Also published as: EP0269642A4; WO1987005362A1

Abstract

Une charge propulsive solide pour un moteur-fusée ou un générateur de gaz comprend un agent propulsif en masse (2) ayant des noyaux longitudinaux de matériau de régulation (3). Les noyaux de régulation offrent une vitesse de combustion supérieure à celle de l'agent propulsif en masse, laquelle diminue avec l'augmentation de la pression, de telle sorte que la vitesse de combustion de la charge propulsive solide décroît également au fur et à mesure que la pression augmente.A solid propellant charge for a rocket engine or gas generator includes a bulk propellant (2) having longitudinal cores of regulating material (3). The control cores provide a higher rate of combustion than that of the bulk propellant, which decreases with increasing pressure, so that the rate of combustion of the solid propellant charge also decreases as as the pressure increases.

Description

ROCKET PROPELLANT CHARGE WITH NEGATIVE PRESSURE EXPONENT

This invention relates to a novel method of making a solid propellant rocket motor or gas generator charge with a negative burning rate pressure exponent (nega¬ tive dr, /dP) . The negative exponent characteristic 5. can be obtained independently of burning rate and other bulk propellant properties allowing impulse, flame temperature, signature, mechanical properties to be chosen as required.

DESCRIPTION OF PRIOR ART

10. Propellant burning rates are usually described as a function of pressure using Vieille's Law,

r_b = aPⁿ (1)

The use of negative pressure exponent propellant (negative n) has been advocated for use with variable 15. area nozzles especially for gas generator applications, see Cohen, J. , Landers, L.C. and Lou, R.L., "Minimum- Response-Delay Controllable Solid-Propellant Gas Generators". J. Spacecraft, Vol. 14, No. 5, May 1977.

A gas generator containing a propellant with 20. a very large negative exponent behaves like a constant pressure source. A few examples of negative exponent propellants have been created as shown in Cohen et al above. However their ballistic properties are not fully understood. In the few examples of negative 5. exponent propellants, the negative exponent character¬ istic is usually gained at the expense of other desir¬ able ballistic properties.

FORCED. CONE BURNING OF A SINGLE CORE

Generally rocket motors may be classified accord-

10. ing to whether they employ liquid fuels, usually com¬ prised of separate reservoirs of oxidizing and oxid- izable materials, or solid fuels, usually comprised of a solid state mixture of oxidizing and oxidizable materials formed into a single fuel grain mass. Liquid

15. fuel rockets depend upon burning which occurs with contact of the oxidizing and oxidizable materials, such burning is initiated by a small source of external heat while solid fuel rockets depend upon burning which is initiated by elevating the surface temperature

20. of the oxidizing-oxidizable mixture in at least a small region to the temperature of ignition. Once ignited, the burning rate of a solid fuel rocket can be increased by preheating or raising the temper¬ ature of small areas in the fuel mass just prior to

25. their participation in the burning reaction. Solid fuel rockets are preferred for military and long mission uses because of their inherent simplicity and avoidance of the complex plumbing, mixing and control elements required in liquid fuel rockets and because 5. of the ease and safety with which the rocket fuel, or grain, can be handled and stored for future use. The design parameters, fuel selection tradeoffs, and operating characteristics of solid fuel rockets are well known in the art and are, for example, discussed 10. in the referenced textbook "Rocket Propulsion Elements", 2d ed., by G.P. Sutton, and "Propellant Chemistry" by Stanley F. Sarver. The disclosure of these referenced texts is hereby incorporated by reference into the present specification.

15. Modulation or termination of the thrust-producing reaction once grain burn has been commenced poses some difficulty in a solid fuel rocket; this difficulty is a characteristic of solid fuel rockets and makes it desirable to employ apparatus such as is described

20. herein to achieve some degree of burn rate modulation or even thrust termination ability. A need for thrust modulation or thrust termination and re-initiation can be readily appreciated in military or scientific rockets. The functions of threat.avoidance, multiple

25. purpose missions and vehicle atmospheric re-entry each present a desirable environment for some form of thrust change, for example.

The present apparatus is principally concerned with thrust modulation or burn rate control during 30. the grain burn of a solid fuel rocket. Such thrust modulation might, for example, also be desirable in tailoring the orbit of a spacecraft, in trading thrust magnitude for thrust duration in a particular rocket application, or in balancing the thrust applied to a parallel rocket vehicle, especially during the initial liftoff, low air velocity, flight portion. Burn rate control can also be useful in achieving fixed levels of thrust which are independent of fuel temperature 5. variations resulting from atmospheric conditions.

The prior patent art includes several examples of arrangements achieving a degree of control over solid fuel rocket propulsion systems and includes the patent of J. Trotel, U.S. 3,457,726 which discloses

10. a plurality of layered fuel arrangements that achieve incremental or intermittent control of the docket thrust by way of periodically enabling the burn of new fuel increments. The Trotel apparatus contemplates the separation of fuel increments by inhibitor layers which

15. are immune to the temperature of rocket operation and which must be violated for renewed fuel access by the use of external energy such as heat or heat-producing electrical energy. The Trotel apparatus relies on these inert fuel-separating layers and their suscepti-

20. bility to external energy as a thrust termination or periodic modulation mechanism and is unconcerned with the effect of external energy on the fuel grain per se^'.

The patent of Ju Chin Chu, U.S. -3,732,693 describes 25. a gel fuel apparatus, wherein fuel of a class intermediate the usual liquid and solid fuel types is utilized by way of pressure feeding, preheating, and use of a granu¬ lated oxidizer in order that controllable rocket thrust be attained.

30. Another example of a solid fuel rocket control apparatus is found in the patent of J.E. Picquendar, U.S. 3,398,537 which also employs externally supplied electrical energy to maintain a supply of combustible fuel to the thrust generating reaction. The Picquendar invention employs a grain composition that is selected to give a non-self sustaining burn; this provides a rocket motor that is responsive to externally applied heat energy. The Picquendar apparatus

5. contemplates the use of radiant heat from a fixed-location resistance heat source and the use of this heat-based control mechanism principally for initiating and stopping thrust generation. The heater of the Picquendar apparatus is so disposed with respect to the grain as to be minimally

10. and inconsistently effective in providing burn surface geometry change.

Another example of prior art solid fuel rocket arrangements is found in the U.S. Patent of Hisao Oka oto et al, U.S. 4,369,710 which discloses an end burn

15. arrangement of solid fuel wherein burn is enhanced through the use of heat conducting filament elements buried in the grain at manufacture and exposed to burn chamber temperatures during rocket operation. The Okamoto invention contemplates the forming of one or more cones

20. in the fuel grain as a result of the preheating achieved with the heat conducting buried filaments. The Okamoto apparatus is principally concerned with the improvements achievable using the fundamental buried filament concept without extension of this concept-into a modulation

25. or control arrangement.

The patent of A.P. Ada son, U.S. 3,065,597 discloses a solid fuel rocket which is capable of the extinguishing and re-ignition functions through the use of burn chamber pressure control. The Adamson invention is based on 30. the concept of burn in the rocket pressure vessel being dependent upon the presence of pressures above a certain threshold for continuation. The Adamson apparatus provides an arrangement for increasing the burn chamber temperature by external means up to the threshold of burn maintenance when re-ignition of the rocket is desired.

5. Another prior art solid fuel rocket control arrange¬ ment is found in U.S. Patent of J.J. De Haye, U.S. 3,529,425 which discloses the use of electrodes which locate a preheating of an electric arc at the grain burn surface or alternately electrodes having temperature-

10. responsive variable electrical resistance which is automatically activated for achieving preheating by the approach of the grain burn surface. The De Haye apparatus also contemplates the achievement of coning at the grain burn face through increased grain burn

15. rate and application of the externally-sourced electrical energy.

Another example of a restartable solid fuel rocket motor is found in the Patent of R.D. Wolcott, U.S. 3,248,875 which describes the use of electrically-heated 20. igniter bands to re-initiate fuel grain burn.

Other examples of solid fuel rocket motor control are found in the patents of R.L. Rod, U.S. 3,066,482 and G.H. Messerly, U.S. 3,182,451 which concern respec¬ tively the achievement of increased burn rate by the 25. addition of acoustic or other transponder-supplied elastic wave (vibratory) energy for increasing fuel combustion efficiency and the use of fluids in conductive tubes which pass through the body of the fuel grain for controlling the temperature of the fuel grain. Additional examples of solid fuel rocket motor burn rate control are to be found in the Patents of R.H. hitesides, JR., U.S. 4,345,427 R.L. Glick, U.S. 3,381,476 and L.H. Caveny, U.S. 3,630,028 which concern 5. improvement in the control apparatus employed with a retractable filament burn rate control variations of the retractable filament structure and the addition of grain cutter elements to the ends of retractable filaments, respectively.

10. It is well known from the article by Summerfield M. and Parker, K.H. entitled "Mechanics and Chemistry of Solid Propellants" Edited by Eringen, A.C. et al pp 75-116. Pergamon Press 1967, and also the articles by Caveny, L.H. and Glick, R.L. "Influence of Embedded

15. Metal Fibres on Solid-Propellant Buring Rate". J. Space¬ craft and Rockets Vol. 4, No. 1 Jan 1969, and Bradfield, W.A. "The Use of Cavities in an End Buring Solid Propellant charge..." Weapons Research Establishment Tech Memo CPD 102, Sept 1968, that an end burning charge

20. with a longitudinal inclusion of a core of higher burn rate will reach an equilibrium surface shape of a cone, of included half angle

_Θ - ^{sin_1 t}- ' ^{2 )} where r, is the burn-rate of the propellant, and r,

25. the (higher) burn rate of the core see Winch, P.C. and Irvine, R.D. "Forced Cone Burning for Active Control of Solid Propellant Burning Rate". AIAA J. Propulsion and Power. Both r,b and r,be can be functions of the motor pressure P, which is a function of the^' propellant 30. burning rate and surface area, the nozzle throat area, and transiently, the rocket motor free volume. BRIEF STATEMENT OF THE INVENTION

The object of the present invention is to provide certain improvements to thrust control of a rocket, and this is achieved according to this invention by 5. using a small quantity of high-negative-exponent pro¬ pellants as core material in charges, the bulk of which are made of conventional propellant materials. The bulk propellant is chosen to provide ballistic properties such as specific impulse, density, smoke, 10. flame temperature or other properties important in a particular application and not necessarily available from the negative exponent propellant.

The charge is constructed with a core, or several cores of propellant with an intrinsically negative

15. pressure exponent with the bulk of the propellant made up of any propellant with a lower burning rate than the core propellant over the pressure range of interest. The core propellant burning rate range can be adjusted to give the desired value by a method

20. of burning rate acceleration.

The result is a charge with pressure and temper¬ ature sensititivies determined by the cores, burning rate determined by the acceleration, and specific impulse and other bulk properties determined by the 25. bulk propellant.

Suitable core materials and burning rate acceler¬ ation methods are referred to herein, and the use of such a charge in a rocket motor or gas generator with a variable area nozzle device to give controllable thrust is included.

The resultant charge has a burning rate and pressure sensitivity determined by the core negative 5. exponent material, but other ballistic characteristics such as specific impulse, flame temperature and smoke are determined by the properties of the bulk material.

The charge design incorporates fibres or strands of core material embedded in the matrix of the bulk 10. propellant and aligned generally perpendicular to the intended burning face. The cross sectional area of the cores can be very small compared to the overall burning surface area, so that the mean properties are essentially those of the bulk propellant.

15. Cone burning is forced around each core by arrang¬ ing that the core material over the operating pressure range of the motor has a higher burning rate than that of the bulk material.

Thus the core regresses faster than the bulk 20. propellant, exposing extra surface area in the bulk propellant in the form of a cone around the core, until the cones from adjacent cores intersect. This quasi-steady-state surface, consisting of a large number of intersecting cones, now regresses at a linear 25. rate equal to the core burn rate. The bulk propellant is consumed at exactly the same rate as if the original planar surface had re¬ gressed at the core burn rate. In other words, the charge produces a gas whose properties are those of 5. the bulk propellant at the mass flow rate of a charge made of the core propellant. A gas generator would now be acting as if it were a constant pressure source, as the quasi-steady-state burning rate has a negative pressure exponent.

TO* Such a charge with a variable area nozzle device gives a positively controllable burning rate. This is used with a high specific impulse bulk propellant in controlled thrust rocket motors. It will also, with a low flame-temperature propellant, provide a control-

15. lable gas supply for long duration gas generator applications.

BRIEF DESCRIPTION OF THE DRAWINGS

In the following description of the invention, reference will be had to the accompanying drawings 20. in which,

FIG. 1 is a schematic central section of a charge showing the bulk propellant, the core, and the inhibitor,

FIG. 2 is a graph of burning rate against pressure 25. showing typical negative exponent propellants reported by Cohen, Landers and Lou, FIG. 3 is a graph similar to FIG. 2 showing the negative exponent Extruded Double Base propellant produced by the applicant,

FIG. 4 is a view similar to FIG. 1 showing the 5. development of burning surface from initial flat end to approximately conical surface by use of an axial line of spaced cavities,

FIG. 5 is a longitudinal section of a negative exponent strand of Bradfield cavitie's,

10. FIGS. 6 and 7 show a negative exponent charge in diagramatic longitudinal and cross section respect¬ ively using strands of Bradfield cavities dispersed as spaced burning control means arranged as a regular pattern through the charge, and

15. FIG. 8 shows the negative exponent charge controlled by means of a variable area nozzle device.

In FIG. 1 the inhibitor 1 surrounds the bulk propellant 2 which in turn has the core 3 axially dispersed within it.

20. DESCRIPTION OF THE PREFERRED EMBODIMENT

EQUILIBRIUM CONICAL SURFACES

Consider a single core fibre of vanishingly small diameter of burning rate r, . This fibre is embedded centrally in a cylinder of propellant of burning rate 25. r_b. The cylindrical matrix has a diameter D and is inhibited from burning on all faces except the end face. The single core charge is illustrated in FIG. 1.

If for any pressure, the bulk burning rate r, 5. is greater than the core burning rate r, , then the charge burns planar and is of no interest. Therefore consideration is restricted to the pressure range where r^r^.

It is well known from Winch and Irvine and others 10. that, if r, is greater than r. , the surface will transform from a flat surface, to a concave cone of semi angle 0, which is the equilibrium surface for the geometry. Thereafter, the surface regresses with a constant cone angle ®, which is a function only 15. of the ratio of the burning rates.

The volume of propellant consumed per unit time is the same as the original flat surface would consume if it were burning at the accelerated burning rate r, , rather than r, . Thus the charge produces gas 20. at a rate which is determined by the core burning rate, but the nature of the gas produced is that of the bulk propellant.

This invention makes use of the fact that, by extension, if the core propellant burns with a negative burning rate pressure exponent, the entire charge produces gas at a rate determined by that negative pressure exponent characteristic.

TRANSITION TIME

5. Control calculations are discussed in Winch and Irvine who have calculated the transition times for a variety of conditions, in units of D/2r, , the time taken to burn from the centre of the cylinder to the edge. The time taken to burn from one cone to another 10. is always at least D/2r, , but for cones steeper that 45°, the time is never greater than .

For many applications, the charge burning rate, set by the core burning rate r, , may be required to vary ^J over a set rang °e r,bc_m,m. to r,b„c_m,a„„x to satisfy ^J charg °e

15. geometry and motor thrust requirements. It can be shown from Winch and Irvine that the worst case transition time is minimised by choosing r, , the bulk propellant burning rate, so that the minimum cone angle θ__in equals

45 or r,bcmm. = j/2r,b. The maximum transition time

20. is then, as above,

" ^D/rbcmin

LIMIT FOR INFINITE NEGATIVE EXPONENT ^'

The limit of infinite negative exponent is a step

25. decrease in burning rate as the pressure rises over a vanishingly small pressure range. The important parameter is the magnitude of the change in burning rate,' expressed as the ratio of r,bcmax to r,bcmm. . Cohen,'

Landers and Lou above refer to this ratio as the tum- 30. down ratio. The resulting charge will, at equilibrium, burn at a constant pressure, regardless of nozzle area. The variable nozzle area device must be appropriate for the burning° rate rang°e r,bcmax to r,bcm. , but within that range the motor will find an equilibrium burning rate and an equilibrium surface area to match the nozzle.

5. In this sense, a negative exponent charge with infinite exponent is identical to a constant pressure source. The chamber pressure is constant, independent of the nozzle area, while the mass low is directly proportional to the nozzle area.

10. A charge with a central core of negative exponent propellant and a bulk of slower burning propellant of different properties, will have a burning rate at equilibrium equal to the negative exponent propellant, and gas properties of the bulk propellant. The time

15. to reach equilibrium can be kept to less than D/r, . , where D is the diameter of the g °rain and r.bcmm. is the minimum burn rate of the core. Although a new dimension is added to the performance and flexibility of charge design, it comes with the penalty of transition

20. times. This transition time, τ to reach equilibrium will be a limiting feature of the charge performance. The transition time can be reduced by reducing the charge diameter, or increasing the minimum core burning rate. There are many. systems limitations on charge

25. diameter, and with so few negative exponent propellants particular burning rates are not available.

MULTIPLE CORE CHARGES

The transition time can be reduced by incorporating many cores of negative exponent propellant, equally 30. spaced across the grain. Winch and Irvine have discussed the use of multiple cores for forced cone burning to minimise the transition time and their results. For multiple cores, the transition time becomes Z/r_b . where Z is the mean distance between cores. If the cores are optimally spaced, Z is approximately equal to D/N where N is the number of fibres.

The significant parameter is the ratio of the transition time τ to the maximum duration of firing

t- = L/r, . f bcmm

where L is the length of the charge. Hence

10, Table 1 lists the ratio τ/t_f for various L/D ratios and various numbers of cores, N. It can be seen, for instance, that to achieve a value of τ/ _{f 0}f less than 0.05 requires an L/D greater than 4, with N greater than 20, while a τ/t_f of less than 0.01 needs high values of L/D,

15. greater than 8, and more than 100 cores.

Table 1. Transition time/Burn time ratio for various L/D and N.

N

0,

The reference to Winch and Irvine above also discusses the useful lower limit to τ/t_f, given by the time constant of the volume of the rocket motor case. There is little value in reducing the transition time to less than 5. this volumetric time constant, which turns out to depend almost, entirely on chamber pressure. If the charge L/D ratio, and the number of cores are sufficiently large, then the charge made using negative exponent core technology will have a thrust response limited 10. only by the response time of the chamber volume to a change in nozzle area.

NEGATIVE EXPONENT PROPELLANTS SUITABLE FOR CORE MATERIAL

Previous negative exponent propellants

Cohen, Landers and Lou report a family of cast 15. composite ammonium perchlorate based propellants, with negative exponents. The significant parameters quoted are the exponent n and the turn-down ratio ^r _maχ ^r _ι5πι-_n*

The propellants quoted vary from cool propellants for gas generator applications with n of -2.7, and

20. a turn-down ratio of 2, through moderate specific impulse propellants with n of -1.0 and turn-down ratio of 2.6, to high specific impulse aluminised propellants with n of -0.5 and turn-down ratio of 1.5.< The minimum burning rate for the gas generator propellant was 1.5

25. mm/s, while for the higher specific impulse propellants, ^rbmin ^{was a}^^out ^~~ ^{mm s}- Typical variation of r, with P is shown in Figure 4.

It has been well known for many years that double base propellants can be modified to give "plateau" 30. (n=0) and "mesa" (n<0) burning over a limited pressure range. Usually the turn-down ratio is quite small, and n is very close to zero over a wide pressure range, but extreme cases are occasionally reported, see Lengelle, G. et al. "Fundamentals of Solid Propellant combustion" 5. Edited Kno, K.K. and Summerfield, M. Chapter 7, Vol.

90 Progress in Astronautic and Aeronautics, AIAA 1984.

Extruded double base propellants produced by the applicant.

A survey of experimental cast double base (CDB) 10. and extruded double based (EDB) propellants made by

Propulsion Division of the applicant yielded the extruded double base propellant with burning rates dependent on pressure shown in Figure 5. as described by Ayres, N. and Odgers, S. "Development of Double Base Negative 15. Exponent Propellants" a WSRL Technical Report in publi¬ cation. This propellant formulation has the following desirable properties:

(i) Turn-down ratio of 2.06.

(ii) Average n over the range r_bmaχ to r_bmin 20. of -2.7.

(iii) Peak n of about -7.

(iv) Minimum burn rate in negative n region of 5.0mm/s.

(v) Very good batch-to-batch and strand-to-

25, strand repeatability.

(vi) Easily producible in "fibre" form.

(vii) Compatible with a wide range of CDB propellants. The maximum and minimum burning rates of this propellant of 10.3 mm/s at 5.5 MPa and 5 mm/s at 7.5 MPa lie in a reasonably useful range, being fairly typical of burning rates used in long burning time 5. end-burning motors.

The turn-down of this propellant is particularly sudden, with the drop in burning rate occurring over such a narrow range of pressure that the strand-burning equipment was unable to hold pressure to a sufficient 10. accuracy to measure the real exponent. The drop from 7-8 mm/s at 7 MPa to 5 mm/s at 7.25 MPa is the limit of resolution of the equipment used in these tests.

Example

The propellant has been extruded in 6 mm, 3 mm and 15. 1 mm diameter strands without difficulty, and no variation in burning rate characteristics has been observed with change in diameter. Strands of the propellant have been incorporated into small charges by the applicant to investigate the practicality of this means of charge 20. manufacture.

Accelerated burning of core material

Winch and Irvine have reviewed a wide range of means of inducing accelerated burning rate in an end- burning charge by forced cone burning. By considering 25. the negative exponent core strand alone, the same ideas of forced cone-burning can be carried over to increasing the core burning rate. It will be seen that if the core strand had, for example, a central wire core as in the invention of B. Silver U.S. Patent 3140663 and the work of Caveny and Glick, the thermal feedback from the flame-zone would increase the burning rate of the core strand and the incorporation of such a wire is optional in this invention.

One means of amplifying the burning rate that 5. is independent of pressure and burning mechanism is the use of cavities as proposed by Bradfield. Bradfield has shown that if a spaced line of cavities is included in a charge, then the average burning rate of the line is amplified by the simple geometrical ratio of the 10. pitch spacing of the cavities, P, to the length of the propellant between them, S.

(r,be/r.b)average = P/S

This is shown in figure 4 where the inhibitor 1, the bulk propellant 2 is again shown but central

15. cavities 4 are indicated. The resultant accelerated mean burning rate causes an approximate cone to form in the charge. Of course the pulsed nature of the cavity line leads to a fluctuating slope on the conical surface but Bradfield has shown that the resultant

20. surface area fluctuations are small if the spacing

^'of the cavities is small compared to the charge radius. The use of a number of cavity lines, especially out of phase with each other, decreases any fluctuations still further.

25. If EDB negative exponent propellant core incorporates

Bradfield cavities then the burning rate, at both r,bcma„x and r-.bcmm. ,' would be increased by ^J the ratio of cavity pitch to propellant length between cavities, P/S. The negative exponent and the turn-down ratio 30. would remain unaffected. This is because the propellant still burns with the negative exponent characteristic, but occassionally the flame-front jumps, with almost no time delay, the distance through the cavity. Thus a distance P is burnt in the time taken to burn a distance of S in the propellant. The incorporation of such 5. cavities is optional in this invention.

In practice, any method which rapidly moves the flame front a distance of P-S produces the desired effect of increasing the burning rate without affecting the core pressure exponent or turn-down ratio. For instance,

10.. the incorporation of "Pyrofuse" Al-Pd exothermic alloying fuse wire, which burns at least ten times as fast as the EDB formulation, in the form of short lengths regularly or randomly spaced, might be expected to amplify the burn rate by the ratio (L+G)/G where L is the length

15. of the "Pyrofuse", and G the mean gap between lengths.

Figure 5 shows a suggested negative exponent strand made using Bradfield cavities to double the burning rate of the EDB propellant to a range 10 mm/s to 20 mm/s. Each strand is constructed of a large number of

20. small cylinders, using solid cylinders 5 alternating with hollow cylinders 6, of equal length, extruded from the same negative exponent propellant, and glued together with an appropriate solvent adhesive. Any amplification is possible by using different lengths

25. of hollow and solid cylinders.

Figures 6 and 7 show a multi-core charge with a core burning rate of 10 mm/s to 20 mm/s, a bulk burning rate of 8 mm/s and a transition time from one burning surface area to another of 0.01 t, , where t, is the 30. time of burn at the lower rate.

APPLICATIONS

The charge illustrated in Figure 6 shows the propellant 2 in a case 8 and shows how a multiplicity of spaced cores 3 are disposed. This assembly has combined several very useful pioperties.

Firstly, it has a large negative exponent and 5. therefore can be used as a controllable gas generator or thurst source with little variation in chamber pressure.

Secondly, the actual burning rate range can be adjusted to give a useful mass flow rate range for a particular application.

10. Thirdly, and without losing the above characteristics, the bulk propellant can be chcsen from a wide range of available propellant formulations to give particular requirements, such as high impulse (aluminised CDB, for example), low smoke (non-alu inised CDB), or low

15. flame temperature, requirements which would vary from application to application.

In a variable are a nczzle device controlled msss flow rocket motor or gas generator the burning rate must be a strong function of pressure for any variation

20. of thrust with nozzle area to occur. In the past, high positive exponents, o<n<l, have been used but they have the disadvantages that small changes in nozzle area cause large changes in pressure.' Thus the motor case must be designed for much higher pressures than

25. the mean pressure. At the same time, with a wide range of pressures to design over, the combustion and nozzle expansion will be far frcm optimum over some of the range. A further disadvantage is that response is slow, with a decrease in nozzle area leading to a decrease in thrust

30. initially, and an increase in thrust after the chamber reaches the new pressure. Cohen, Landers and Lou proposed and demostrated that using negative exponent propellant charges gives a much narrower range of chamber pressures, which optimises both motor case and propellant combustion design, as 5. well as giving much better control response rate.

Control response rate is improved because opening the nozzle instantly gives more thrust, and this is followed by an increase in burning rate to sustain the higher level of thrust with a higher rate of gas production.

10. The negative exponent cored charge described in this invention has most of the advantages of a charge made entirely of negative exponent propellant. The pressure remains nearly constant, and thus the motor chamber and nozzle can be designed around that constant

15. pressure, with only a small safety factor. However, the speed of control does not reach the very rapid response of the pure negative exponent charge and so the control of thrust with the variable nozzle must be considered fully to ensure stable operation. The

20. response is still much faster than the positive exponent alternative, which is slowed down by the need to fill the chamber to a considerably different pressure.

The burning rate range over which such a charge can be operated is extremely wide. Bradfield has demonstrated

25. amplifications of 6, and factors of 10 seem reasonably practical. This could give charge burning rates up to 100 mm/s, if required. The lower limit of burning rate with the negative exponent propellant reported here is 5 mm/s, but Cohen, Landers and Lou report a

30. cast composite propellant with a turn-down ratio of over 2 and a minimum burning rate of 1.5 mm/s. Thus the means are available to make negative exponent cored charges with turn-down ratios of at least 2, over a range of minimum burning rates of 1.5 mm/s to at least 50 mm/s. Although Figure 6 shows an end-burning charge In the form of a cylinder, the invention also applies to any other geometric arrangement of solid propellant in which cores of negative exponent propellant are arranged to burn faster 5. than the bulk propellant, and in so doing expose additional bulk propellant surface area. Such an alternative arrangement might, for instance, be a charge with a central perforation, and the cores arranged radially.

ADVANTAGES

10. The great advantage of the negative exponent cored charge described here is the separation of the ballistics of charge design for controlled nozzle solid propellant motors intp three separate areas:

(a) The negative exponent core material

15. (b) The burning rate of the core

(c) The bulk properties of the charge

Each area can be independantly optimised, so that the resultant charge is no longer such a compromise, and the range of propellants available to the designer 20. is enormously increased.

Initially a designer would select a propellant for the core to have a suitable high negative exponent, and turn-down ratio. Values of n of less than -2.5 have been produced in EDB strands, with a turn-down ratio 25. of greater than 2. Cast composite propellants are also available, with n less than -1.0 and turn-down ratios greater than 2.0 (and up to 2.6).

The designer, having selected the core propellant, can then choose an accelerator (such as Bradfield cavities, or similar devices that cause the flame-front to progress very rapidly through a fixed fraction of the propellant) to raise the burning rate range of the core propellant to the range required by the application.

5. Finally, the designer is free to choose the bulk propellant from any of the wide range of available propellants with suitable propellant compatibility with the core propellant. A minimum bulk burning rate of a third of the minimum core burning rate up to a

10. maximum equal to the minimum core burning rate would probably be acceptable. The designer can design the bulk for maximum specific impulse, low plume signature, maximum volumetric specific impulse, low cost or other properties, while ignoring pressure exponent (except

15. to avoid values near unity for safety reasons) and temperature coefficient.

ADDED APPLICATIONS

Gas generators for pneumatic actuators are an application where controllability is the most important 20. feature. Rapid response rate, and the ability to choose low flame temperature propellants are particularly attractive.

Such a charge could also be used in gas generator fuelled ram-jets or ducted rockets. In this application,

25. a very fuel rich propellant is burnt in a chamber, exhausted into a duct where it is mixed with air, and the resultant mixture is burnt and exhausted as a ram¬ jet. The control is used to give varying mass flow to compensate for varying air flow at different flight

30. conditions. In this application, speed of response can be quite slow and turn-down ratios of 2 may be adequate, but the ability to choose a very fuel rich propellant without compromising pressure exponent or burning rate is extremely important. The invention can be used in the hovering rocket area which imposes a need for varying thrust rocket motors. Conventional controllable rocket motors pay considerable penalties for that controllability. A 5. negative exponent cored charge with a controlled area nozzle could give thrust control over the limited range required. A turn-down ratio of 2 would be quite satisfactory, while a response time of 1/lOOth of the burn time is achievable and would probably be satisfactory 10. for height control. High specific impulse in the bulk propellant would be a large gain in a vehicle where the thrust is required to support the weight.

Lastly, a controlled thrust rocket motor could in theory replace many of the boost-sustain and pulse-

15. boost motors presently in use. For these applications, turn-down ratios of up to about 5 would be required, and these are not yet available. Further research into negative exponent propellants is needed before this use could be seriously addressed. Response time

20. is not very important, but high specific impulse and low signature in the bulk propellant are high priorities.

CONCLUSION

A method of designing an end-burning solid propellant rocket motor charge with a large negative effective

25. pressure exponent has been presented. This method uses cores of negative exponent propellant embedded longitudinally in a charge of conventional propellant. This method of charge design has the significant advantage that it separates the problems of charge design into

30. three independent areas. (i) A negative exponent core propellant material, which provides the negative exponent property to the whole charge. Several examples of such propellants have been presented.

5. (ii) A burning rate amplification method, which increases the burning rate of the core propellant material to give any desired core burning rate, without affecting the negative exponent. Possible methods of burning rate amplification have been proposed.

10. (iii) A bulk propellant, which can be chosen to give optimum properties for the application intended without affecting the charge burning rate or negative exponent.

The resultant charge has all the advantages of 15^".. negative exponent charges produced entirely from negative exponent propellant, except for the almost instantaneous response of the completely homogeneous negative exponent charge. The negative exponent core charge can, however, be made with as small a response time as required, provided 20. sufficiently many closely spaced cores can be embedded in the charge. The cored charge has none of the major disadvantages of the homogeneous negative exponent charge, because the bulk propellant can be selected to give whatever performance is required.

Claims

THE CLAIMS DEFINING THE INVENTION ARE AS FOLLOWS:-

1. A rocket motor solid propellant charge or gas generator solid propellant charge comprising a body of arbitrarily selected bulk propellant (2) con¬ taining a longitudinal core or cores of control ma er¬ s' ial (3) characterized by said control material having a burning rate, greater than that of the bulk propel¬ lant, which decreases as pressure increases over some range of pressure or in other words a negative burning rate pressure exponent whereby said solid propellant i0. charge has a burning rate which also decreases as pressure increases over said range of pressure.

2. A charge as claimed in claim 1 wherein the said control material (3) is in the form of a series of parallel cores uniformly or otherwise spaced longi¬ tudinally extending from an intended burning surface

5* through the body of bulk propellant (2).

3. A charge as claimed in claim 1 and claim 2 wherein the core or cores of control material is in the form of axially spaced sections (5) separated from each other whereby the burning rate of the core

5- can be varied from that of the control material.

4. A charge as claimed in claim 3 wherein the sections (5) are separated by intermediate sections (6) having a different burning rate.

5. A charge as claimed in claim 3 wherein the sections are alternately solid (5) and hollow (6).

6. A charge as claimed in claim 3, 4 or 5 where¬ in a heat conductive member is included in the said core.

7. A rocket motor or gas generator comprising a charge as claimed in claim 1 combined with a variable area nozzle or valve device whereby mass flow rate from the rocket motor or gas generator can be controlled.

8. ^*A rocket motor or gas generator as in claim

7 wherein the variable area nozzle or valve device (9) consists of a number of nozzles or valves which can be independently opened or closed.

9. A gas generator as claimed in claims 7 and

8 wherein the bulk propellant is selected to burn to produce fuel-rich combustion products suitable for combustion in a ducted rocket motor whereby the mass flow rate of fuel-rich combustion products into the secondary combustor of said ducted rocket is controlled.

10. A gas generator as claimed in claims 7 and 8 whereby a constant pressure is maintained in the gas generator chamber independently of mass flow rate by using variable area nozzle or valve devices.

11. A rocket motor as claimed in claims 7 and 8 whereby the magnitude and direction of the thrust is controlled.

12. A rocket motor or gas generator as claimed in claims 7 and 8 combined with a temperature sensor and nozzle area controller whereby the thrust or mass flow rate is independent of ambient temperature.