US2804824A - Turbine speed regulators - Google Patents

Description

Sept. 3, 1957 OR. CRUZAN 2,804,824

TURBINE SPEED REGULATORS Filed May 13, 1954 INVEXTOR Orvul R Cruzan WEMMM W ATTORNEYS 2,804,824 1 Patented Sept. 3, 1957 TURBINE SPEED REGULATORS Orval R. Cruzan, Washington, D. C., assignor to the United States of America as represented by the Secretary of the Army Application May 13, i954, Serial No. 429,709

2 Claims. or. 102-701 (Granted under Title 35, U. S. Code (1952), sec. 266) The invention described herein may be manufactured and used by or for the Government for governmental purposes without the payment to me of any royalties thereon.

This invention relates to air driven turbines for driving electric generators in missile fuzes and more particularly to regulating means for air driven turbines in fuzes of supersonic missiles.

One object of the invention is a new and novel regulatory means for turbo-generators used in electronic or electric fuz es of supersonic missiles.

Another object of the invention is a novel regulatory means requiring no moving parts for electronic and electric fuzes of supersonic missiles.

A further object of the invention is static regulatory means for a turbo-electric generator used in proximity fuzes of supersonic missiles wherein a substantially constant voltage is obtained at the output of the generator regardless of the missile speed variations in the supersonic range.

A still further object of the invention is a turbo-generator regulating means for supersonic missiles which is economical of manufacture and which readily lends itself to mass production.

The specific nature of the invention as well as other objects and advantages thereof will clearly appear from the following description and drawing wherein:

Figure 1 is a longitudinal sectional view of a fuze embodying this invention taken along line 11 of Fig. 2.

Figure 2 is a cross sectional view of the fuze shown in Fig. 1 taken along line 22.

Figure 3 is a perspective view of a supersonic missile in which is incorporated a fuze employing the invention.

As is known, it is desirable to maintain a substantially constant output voltage at the generator terminals during the active period of the fuze. this requires the turbo-generator to operate at substantially constant speed during the said active period.

The invention comprises a fuze for supersonic missiles having formed in the base thereof a chamber containing a turbine rotor and provided with intake and exhaust ducts, the ratio of the cross sectional area of the intake duct to exhaust ducts being such that the velocity of the turbine is maintained constant at supersonic speeds.

In the supersonic range if the pressure at the exhaust ports is maintained substantially constant, and the intake 60 pressure is increased, the speed of the turbine rotor increases, the speed of the rotor being proportional to the velocity of the air past the rotor blades or vanes which is proportional, in turn, to the velocity of the air leaving the exhaust ports. Eventually, however, in the supersonic range a pressure diflerential is reached beyond which any additional pressure differential will not be reflected in an increase in the velocity of the air from the exhaust ports. The reason for this leveling ofl effect is that a gas, under pressure, flows no faster than the velocity of sound through that gas under the prevailing temperature conditions.

ture.

One method of doing A fuze of the class described aifixed to a supersonic missile will experience a suflicient nose pressure to maintain the air through the exhaust ports atsonic speed. Inasmuch as the air through these ports cannot increase 5 beyond that speed, additional missile speed will not cause an increase in the air velocity past the vanes. Consequently, the turbine rotor speed will be maintained constant provided the missile is traveling at sonic speed or above. I r

The previous discussion is predicated upon isothermal -pressure variations. The variations in turbine chamber pressures under flight conditions will be adiabatic, however. Consequently, an increase in nose pressure of the fuze will be accompanied by an increase in air tempera- Inasmuch as the velocity of sound is proportional -,-.to the temperature of the sound conducting medium, an increase in missile speed, even in the supersonic range, is accompanied by some increase of air velocity past the turbine rotor vanes. At low'supersonic velocities the eifect is only slight.

A preferred embodiment of the-invention is shown in Figs. 1 and 2 wherein the reference 1 generally indicates the fuze having abase 10 which is closed at its forward endby cover plate 11. The cover plate is secured to the through a plurality of apertures 13 in the periphery of shoulder 14. The duct 15 has threaded portion 16 at one end thereof which is affixed to threaded portion 17 of aperture 18 located centrally of cover plate 11. The other end of duct 15 has an enlarged portion 19, the opening therein comprising the intake port 19a. Enlarged portion 19 of duct 15 is provided with shoulder 20. A thin walled ogive 21 is positioned between shoulder 20 of r duct 15 and shoulder 22 of cover plate 11. The ogive 9 21 serves to support the duct 15 as well as to inclose and protect electronic components, not shown, which are contained within the ogive. Located below cover plate 11 and within turbine rotor chamber 33, formed in the forward part of base 10, is turbine rotor 24. Turbine rotor 24 is mounted upon and adapted to turn with shaft 25, which shaft is journaled within support plates 26 and 27. Support plate 26 rests upon shoulder 28 of base 10. Turbine rotor 24 is provided with a plurality of vanes 29, Fig. 2. Within base 10 and resting between support plates 26 and 27 is generator 31 which is adapted to turn with shaft 25. Support plate 27 rests upon shoulder 32 of base 10. The rearward portion of base 10 contains the detonator and booster assembly, not shown, of the fuze. Radially disposed exhaust ducts 34 are formed in the Wall 10a of base 16 and are in communication with turbine rotor chamber 33 and flared exhaust ports 34a in the peripheral surface of wall 10a. Safety pin 35, Fig. 2, is adapted to be inserted into or extracted from aperture 36 in base 10. When in place within the fuze safety pin 35 engages notch 37 in turbine rotor 24 to prevent the rotor from prematurely turning.

The invention functions in the following manner: As the missile with the fuze employing this invention leaves the launcher and gains speed as it moves through the air, a nose pressure builds up in front of the fuze which is greater than the pressure at the exhaust ports 34. Consequently, air is forced into the duct 15 through port 19a, past the turbine vanes 29, through ducts 34, and out the exhaust ports 34a. As the air passes the said vanes the said turbine rotor is caused to spin. The electric generator 31 is adapted to turn with the turbine rotor upon shaft 25. As the speed of the missile increases the pressure diflerential between the turbine intake and the exhaust ports increases, causing the air velocity past the base 10 by a plurality of screws 12 which enter base 10 i of the missile in the supersonic range causes anincreased pressure differential but not'an. increase in the air velocity a past thelturbine blades'. Consequently,lthere is no increase .of speed of the turbine rotor. It follows that any fluctuations in missileyelocity inthe supersonic range will" not be reflected inhthe turbine speed. 7 Thus, the tur-V bine' speed will be" maintained substantially constant. This is predicated upon. the principle that a gas will'not flow at 'avel-ocity exceeding thespeed of sound through that gas,:rega1;dle'sslof the amount of .pressure to which the.gas-is subjected. The result'is that the electric generator'of the fuze'is driven at constant speedduring anyperiodwhenthe supersonic missile 38 has attained supersonic speed.

It will'be apparent that the embodiment shown is only exemplary and. that various modifications can be made i in construction and arrangement within the scope of .the'

inventions as definedin the appended claims I claim:

. 1. "In 'an ordnance missile'having an air turbine driven by' air' flow: resultingfromxmissile flight, an air-supply structure adaptedto substantiallylimit the maximum speed attainable by'said turbine; said structure compri sing: a housing general-1y enclosing said .turbine; an air;

intake structure for admitting air to said housing, said intake-structure-having at leastone'inlet-aperture facing generally forward with respect to the axis of said missile, forward flight of said missile being adapted to force air through said inlet aperture andthrough said intake struc ture; an air outlet structure adapted to permit air to flow out of said housing during forwardfiight ofsaid missile,

r the effective minimum cross-sectional area of said outlet structuretbeingflessithah the efiective minimum cross- 7 sectional a'rea-of said intakestructure, the maximum rate of flow of air through said. housing as missilevelocity increases; being limitedby-the circumstance that the maximum velocity attainable by air flowing through said outlet aperture is substantially-equal to the velocity of sound, and vthe maximum speed of said turbine being. correspondin'gly limited.

2. The invention according to claim 1 in combination with an electrical generator i driven: by said turbine, .the

output -of-;said. generator thus .being relatively .independ ent of missile velocity-provided-missile:velocity is at least.

suflicient to cause air to flow through said outlet struc-. ture at substantially thespeedofi sound. 4

References Cited-iii thej file of this patent UNITED STATES PATENTS 2,442;7-83-; Senn June-8, 1948 FQREIGN' PATENTS V 304,254.v Germany; .Oct. 1, 192

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