US3004734A - Hydraulic power supply - Google Patents
Hydraulic power supply Download PDFInfo
- Publication number
- US3004734A US3004734A US796181A US79618159A US3004734A US 3004734 A US3004734 A US 3004734A US 796181 A US796181 A US 796181A US 79618159 A US79618159 A US 79618159A US 3004734 A US3004734 A US 3004734A
- Authority
- US
- United States
- Prior art keywords
- power supply
- hydraulic power
- missile
- hydraulic
- rocket
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired - Lifetime
Links
- 239000012530 fluid Substances 0.000 description 13
- 239000007789 gas Substances 0.000 description 11
- 230000001141 propulsive effect Effects 0.000 description 6
- 238000012360 testing method Methods 0.000 description 4
- 230000007797 corrosion Effects 0.000 description 2
- 238000005260 corrosion Methods 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 241000239290 Araneae Species 0.000 description 1
- OKTJSMMVPCPJKN-UHFFFAOYSA-N Carbon Chemical compound [C] OKTJSMMVPCPJKN-UHFFFAOYSA-N 0.000 description 1
- 229920000459 Nitrile rubber Polymers 0.000 description 1
- BQCADISMDOOEFD-UHFFFAOYSA-N Silver Chemical compound [Ag] BQCADISMDOOEFD-UHFFFAOYSA-N 0.000 description 1
- 229910000831 Steel Inorganic materials 0.000 description 1
- 230000004075 alteration Effects 0.000 description 1
- 239000000919 ceramic Substances 0.000 description 1
- 238000004891 communication Methods 0.000 description 1
- 230000000994 depressogenic effect Effects 0.000 description 1
- 229920001971 elastomer Polymers 0.000 description 1
- 230000003628 erosive effect Effects 0.000 description 1
- 239000011521 glass Substances 0.000 description 1
- 229910002804 graphite Inorganic materials 0.000 description 1
- 239000010439 graphite Substances 0.000 description 1
- 239000003779 heat-resistant material Substances 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000012774 insulation material Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000013021 overheating Methods 0.000 description 1
- 229920001084 poly(chloroprene) Polymers 0.000 description 1
- 239000003380 propellant Substances 0.000 description 1
- 229910052709 silver Inorganic materials 0.000 description 1
- 239000004332 silver Substances 0.000 description 1
- 239000004449 solid propellant Substances 0.000 description 1
- 239000010959 steel Substances 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/421—Non-solar power generation
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/40—Arrangements or adaptations of propulsion systems
- B64G1/403—Solid propellant rocket engines
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
- F02K9/32—Constructional parts; Details not otherwise provided for
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F42—AMMUNITION; BLASTING
- F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
- F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
- F42B10/60—Steering arrangements
- F42B10/66—Steering by varying intensity or direction of thrust
- F42B10/663—Steering by varying intensity or direction of thrust using a plurality of transversally acting auxiliary nozzles, which are opened or closed by valves
Definitions
- This invention is concerned with an hydraulic power supply and is more particularly concerned with a new and improved power supply for operating the hydraulically controlled actuators in a missile or rocket or the like.
- the missile is usually steered by fins which are depressed or elevated in a certain manner to stabilize the missile in roll, pitch and azimuth.
- Thrust vectoring is accomplished by means of vanes disposed in the stream of propulsive gases discharged from the rocket motor, by swivel nozzles, or by jetavators.
- the fins, vanes, jetavators or swivel nozzles are powered by an hydraulic actuator or a D.C. motor which operates in response to input signals generated by the guidance section of the missile.
- FIG. 1 is a simplified drawing of the rearward portion of a missile having a hydraulic power supply embodying the principles of this invention
- FIG. 2 is a horizontal section of the hydraulic power supply of this invention
- FIG. 3 is a section taken along line 3--3 of FIG. 2;
- FIG. 4 is a section taken along line 44 of FIG. 2.
- FIG. 1 wherein the hydraulic power supply embodying principles of this invention is shown in a typical missile having a jetavator thrust vectoring system.
- a typical missile having a jetavator thrust vectoring system.
- the particular type of thrust vectoring system or aerodynamic control system is a matter of choice and forms no part of this invention per se.
- Disposed within the missile body 11 is a rocket liner 12 containing a solid propellant rocket motor 13 therein.
- An internal bulkhead 14, bowed slightly away from rocket motor 13, is formed within the missile casing 11 and contains a plurality of thrust nozzles 15 which are in communication with the rocket motor 13 for producing the necessary propulsive thrust.
- a jetavator 16 operated by one of a plurality of hydraulic actuators 17.
- the bulkhead 14 has a central aperture formed therein to receive the forward portion of the housing 18 of the hydraulic power supply unit.
- This housing contains the hydraulic fluid .which is pumped to actuators 17 via lines 19. The return flow from the actuator is discharged throughline 21 as described in greater detail hereinafter.
- Housing 18 is mounted to bulkhead 14 by means of a plurality of bolts 22 which pass through thelmounting flange 23- formed on the housing and thread into appropriate holes in bulkhead 1 4. Threaded into the open end of housing 18 is a baffle plate 44 (FIG. 2) having a plurality of bleed ports 26 therein.
- This baflle plate may be constructed of any'material such for example as 10/20 steel.
- a rather thick slab 27 of heat and corrosion insulation material is bonded to the baflle plate to prevent rapid overheating of plate or erosion thereof by the rocket motor gases which maybe of a temperature up to about 5,000 F.
- Slab 27 may be constructed of graphite, ceramic, or heat resistant glass. It will be noted that the slab 27 also has a plurality of bleed ports 25 which are in alignment with ports '26 in baffle plate 44.
- an O ring 28 made of conventional material such as Neoprene or Buna-N may be used to seal the housing to the bulkhead 14.
- a quantity of hydraulic fluid partially fills the interior of casing 18 as indicated at 29 while a free piston 31 is disposed within the casing, normally adjacent the baffle plate 44- for actuation by the hot rocket motor gases as they bleed through ports 26 against the piston. Since the hydraulic actuator is in operation only for a very short period, it has been found unnecessary to fabricate piston 31 from special high heat resistant material; and O ring 32 like 0 ring 28 may be constructed of conventional rubber.
- the piston 31 is constructed in a manner contrary to the generally accepted practice; that is, the [/11 ratio is much smaller than normal design practice. This permits the use of a casing of smaller exterior size containing a rather large interior volume of fluid 29. That is to say, the low length/ diameter ratio permits a proportionately large volume of fluid to be contained within a given housing. In tests, the piston did not cock or bind as it moved to the right as seen in FIG. 2 under the influence of the high pressure rocket gases. This result is contrary to what would normally be predicted.
- a stepped bore 33 is formed at the end of casing 18 opposite from the baflie plate 44. Disposed within this bore is a spring loaded valve 34 having an 0 ring 35 sealingly disposed against the smaller diameter of bore 33 and supported by spider 24. When valve 34 is moved to the right by pressurized fluid 29, and the seal 35 clears the smaller diameter portion of bore 33, high pressure fluid is allowed to flow by this seal and into passageways 36.
- Passageways 36 run radially out from a bore 33 at intervals and deliver high pressure fluid to one of the servo actuators 17 shown in FIG. 1. :For the sake of convenience, and simplicity these actuators are not shown in FIGS. 2-4.
- An auxiliary port '37 FIG.'3) is also connected to the bore 34 and terminates in a self-sealing connector 38 which may be any available commercial type suitable for the purpose such as a Roylyn connector.”
- Port 37 and valve 38 make it possible to check the operationof the servo valve actuators by applying a high pressure fluid through port 37 and out channels 36 without necessitating the actuation of piston 31 by the rocket inotonthereby making-it possibleito' test their'nissile control section periodically during storage.
- the self-sealing connector 38 prevents loss of poppet valve' 40 (FIG. 1) to prevent accidental loss of fluid prior to ignition of motor 12.
- an hydraulic power supply comprising; a rocket motor for generating propulsive gases, an open ended casing disposed against one end of said motor to receive a portion of the propulsive gases and partially filled with hydraulic fluid, a baffle plate closing theopen end of'said'casing, and having a 1 a plurality of bleed holes to admit a portion of the propulsive gases into the casing, heat and corrosion resistant shield means bonded to said plate to protect it from eroconnected to said channel means for discarding spent 2.
- the hydraulic power supply of claim lfurther comprising; a normally closed port accessible from without the power supply to permit injections of a quantity of testing pressurized fluid thereby providing for periodic-testing of said hydraulic actuators in storage without actuationof said piston.
Description
'UHVW ncrn nuw-u. mmnufl NUUM Oct, 17 1961 J. E, RADFORD HYDRAULIC POWER SUPPLY 5 Sheets-Sheet 1 Filed Feb. 27. 1959 n... "ma.
llllllllulu'nmiifih INVENTOR. JAMES E. RADFORD Adz J. E. RADFORD HYDRAULIC POWER SUPPLY Oct. 17, 1961 3 Sheets-Sheet 3 Filed Feb. 27. 1959 FIG.3.
3,004,734 Patented Oct. 17, 1961 fire 3,004,734 HYDRAULIC POWER SUPPLY James E. Radford, Silver Spring, Md., assignor to the United States of America as represented by the Secretary of the Navy 7 Filed Feb. 27, 1959, Ser. No. 796,181
2 Claims. (Cl. 244-14) (Granted under Title 35, US. Code (1952), see. 266) The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.
This invention is concerned with an hydraulic power supply and is more particularly concerned with a new and improved power supply for operating the hydraulically controlled actuators in a missile or rocket or the like.
It is common practice in the missile art to steer-a guided missile by means of thrust vectoring or by aerodynamic controls. In the latter case, the missile is usually steered by fins which are depressed or elevated in a certain manner to stabilize the missile in roll, pitch and azimuth. Thrust vectoring is accomplished by means of vanes disposed in the stream of propulsive gases discharged from the rocket motor, by swivel nozzles, or by jetavators. Generally, the fins, vanes, jetavators or swivel nozzles are powered by an hydraulic actuator or a D.C. motor which operates in response to input signals generated by the guidance section of the missile. In the past, it has been the practice to power the electric motors by hydraulic powered alternator-s, and to power the alternators by means of high pressure fluid. Both of these systems require an auxiliary hydraulic power unit for generating the necessary high pressure fluid. In the past, this has been accomplished by a motor-pump arrangement wherein the motor is powered by an electrical power supply or by a turbine-pump arrangement powered by high energy gases such for example as those generated by a cartridge containing a propellant.
Both of these systems add to the complexity of an already complex missile and tend to slightly decrease reliability since there is always a danger of malfunctioning of the turbine or the motor which drives the pump.
It is therefore an object of this invention to provide a new and improved auxiliary power supply for a rocket propelled vehicle which power supply derives its energy from the main propulsive gases of the rocket motor.
It is another object of this invention to provide a simple eflicient and rugged hydraulic power supply suitable for use in a rocket or missile.
These and many other objects will become more readily apparent to those skilled in the art when the following specification is read and considered along with the attendant drawings wherein like numerals designate like or similar parts throughout the various views and in which:
'FIG. 1 is a simplified drawing of the rearward portion of a missile having a hydraulic power supply embodying the principles of this invention;
FIG. 2 is a horizontal section of the hydraulic power supply of this invention;
FIG. 3 is a section taken along line 3--3 of FIG. 2; and
FIG. 4 is a section taken along line 44 of FIG. 2.
Referring now with greater particularity to FIG. 1 wherein the hydraulic power supply embodying principles of this invention is shown in a typical missile having a jetavator thrust vectoring system. It is to be understood, of course, that the particular type of thrust vectoring system or aerodynamic control system is a matter of choice and forms no part of this invention per se. Disposed within the missile body 11 is a rocket liner 12 containing a solid propellant rocket motor 13 therein. An internal bulkhead 14, bowed slightly away from rocket motor 13, is formed within the missile casing 11 and contains a plurality of thrust nozzles 15 which are in communication with the rocket motor 13 for producing the necessary propulsive thrust. At the ends of each of these nozzles is a jetavator 16 operated by one of a plurality of hydraulic actuators 17. It will be noted that the bulkhead 14 has a central aperture formed therein to receive the forward portion of the housing 18 of the hydraulic power supply unit. This housing contains the hydraulic fluid .which is pumped to actuators 17 via lines 19. The return flow from the actuator is discharged throughline 21 as described in greater detail hereinafter. i
Since there is relatively little flow of gases past the outside edge of the housing 18 an O ring 28 made of conventional material such as Neoprene or Buna-N may be used to seal the housing to the bulkhead 14.
A quantity of hydraulic fluid partially fills the interior of casing 18 as indicated at 29 while a free piston 31 is disposed within the casing, normally adjacent the baffle plate 44- for actuation by the hot rocket motor gases as they bleed through ports 26 against the piston. Since the hydraulic actuator is in operation only for a very short period, it has been found unnecessary to fabricate piston 31 from special high heat resistant material; and O ring 32 like 0 ring 28 may be constructed of conventional rubber.
It should be noted that the piston 31 is constructed in a manner contrary to the generally accepted practice; that is, the [/11 ratio is much smaller than normal design practice. This permits the use of a casing of smaller exterior size containing a rather large interior volume of fluid 29. That is to say, the low length/ diameter ratio permits a proportionately large volume of fluid to be contained within a given housing. In tests, the piston did not cock or bind as it moved to the right as seen in FIG. 2 under the influence of the high pressure rocket gases. This result is contrary to what would normally be predicted.
A stepped bore 33 is formed at the end of casing 18 opposite from the baflie plate 44. Disposed within this bore is a spring loaded valve 34 having an 0 ring 35 sealingly disposed against the smaller diameter of bore 33 and supported by spider 24. When valve 34 is moved to the right by pressurized fluid 29, and the seal 35 clears the smaller diameter portion of bore 33, high pressure fluid is allowed to flow by this seal and into passageways 36.
By thus formingthe pressure delivery ports and the ex haust-ports in two distinct layers as it were one above the other-in the wall of casing 18, it is possible to conserve space without sacrifice of ruggedness' I It should be apparent to those skilled in the art 'upon reading this specification that this invention provides apparatus for powering the auxiliary power supply unit of a missile directly from the hot exhaust gases of the rocket motor while achieving simplicity of design and dependability of operation. Y
Having thus described this invention with reference to 'but a single preferred embodiment, it is to be understood that this invention may be practiced otherwise than as specifically delineated and is susceptible of many alterations and modifications without departing from the spirit and scope thereof. Accordingly, this invention is not to be construed as limited only by the terms of the appended claims.
What is claimed as new and desired to be secured by Letters Patent of the UnitedStates is:
1. In a missile having a plurality of hydraulic actuator for effecting control of the missile in flight, an hydraulic power supply comprising; a rocket motor for generating propulsive gases, an open ended casing disposed against one end of said motor to receive a portion of the propulsive gases and partially filled with hydraulic fluid, a baffle plate closing theopen end of'said'casing, and having a 1 a plurality of bleed holes to admit a portion of the propulsive gases into the casing, heat and corrosion resistant shield means bonded to said plate to protect it from eroconnected to said channel means for discarding spent 2. The hydraulic power supply of claim lfurther comprising; a normally closed port accessible from without the power supply to permit injections of a quantity of testing pressurized fluid thereby providing for periodic-testing of said hydraulic actuators in storage without actuationof said piston.
References Cited in the file of this patent UNITED STATES PATENTS 2,723,528 Stark Nov. 15, 195.5
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US796181A US3004734A (en) | 1959-02-27 | 1959-02-27 | Hydraulic power supply |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US796181A US3004734A (en) | 1959-02-27 | 1959-02-27 | Hydraulic power supply |
Publications (1)
Publication Number | Publication Date |
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US3004734A true US3004734A (en) | 1961-10-17 |
Family
ID=25167539
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US796181A Expired - Lifetime US3004734A (en) | 1959-02-27 | 1959-02-27 | Hydraulic power supply |
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US (1) | US3004734A (en) |
Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3099959A (en) * | 1961-11-07 | 1963-08-06 | Charles F Bowersett | Rocket engine |
US3136250A (en) * | 1962-05-04 | 1964-06-09 | Samuel A Humphrey | Integrated auxiliary power unit |
US3160098A (en) * | 1962-11-05 | 1964-12-08 | William A Schulze | Missile separation system |
US3167017A (en) * | 1960-11-07 | 1965-01-26 | Gen Motors Corp | Atitude control |
US3333789A (en) * | 1963-07-18 | 1967-08-01 | Jr Charles W Schreiner | Off-on type missile control system |
US3410505A (en) * | 1958-10-17 | 1968-11-12 | Gildon Walter James | Control systems for aerial missiles and like vehicles |
US4599044A (en) * | 1985-01-07 | 1986-07-08 | The United States Of America As Represented By The Secretary Of The Navy | Electronic feedback area control system for TVC gas generator |
US6402091B1 (en) * | 2000-04-03 | 2002-06-11 | Aerojet-General Corporation | Flow-through thrust takeout apparatus |
Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2723528A (en) * | 1955-11-15 | Auxiliary power package |
-
1959
- 1959-02-27 US US796181A patent/US3004734A/en not_active Expired - Lifetime
Patent Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2723528A (en) * | 1955-11-15 | Auxiliary power package |
Cited By (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3410505A (en) * | 1958-10-17 | 1968-11-12 | Gildon Walter James | Control systems for aerial missiles and like vehicles |
US3167017A (en) * | 1960-11-07 | 1965-01-26 | Gen Motors Corp | Atitude control |
US3099959A (en) * | 1961-11-07 | 1963-08-06 | Charles F Bowersett | Rocket engine |
US3136250A (en) * | 1962-05-04 | 1964-06-09 | Samuel A Humphrey | Integrated auxiliary power unit |
US3160098A (en) * | 1962-11-05 | 1964-12-08 | William A Schulze | Missile separation system |
US3333789A (en) * | 1963-07-18 | 1967-08-01 | Jr Charles W Schreiner | Off-on type missile control system |
US4599044A (en) * | 1985-01-07 | 1986-07-08 | The United States Of America As Represented By The Secretary Of The Navy | Electronic feedback area control system for TVC gas generator |
US6402091B1 (en) * | 2000-04-03 | 2002-06-11 | Aerojet-General Corporation | Flow-through thrust takeout apparatus |
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