CN106134428B - Spacecraft rocket engine electric igniter - Google Patents
Spacecraft rocket engine electric igniterInfo
- Publication number
- CN106134428B CN106134428B CN200610120358.3A CN200610120358A CN106134428B CN 106134428 B CN106134428 B CN 106134428B CN 200610120358 A CN200610120358 A CN 200610120358A CN 106134428 B CN106134428 B CN 106134428B
- Authority
- CN
- China
- Prior art keywords
- spark plug
- electric igniter
- voltage
- electric
- spark
- 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 - Fee Related
Links
Landscapes
- Ignition Installations For Internal Combustion Engines (AREA)
Abstract
The invention discloses a kind of spacecraft rocket motor ignition device, comprise that spark plug is connected with driving source by shielded cable, driving source is by oscillating circuit (4), step-up transformer (5), rectification, energy storage, discharge circuit (6) and delay circuit (7) composition; This electric igniter charges to storage capacitor obtaining about 2000VDC voltage after ac voltage rectifier secondary step-up transformer, and in the time that charging reaches certain voltage value, electric capacity produces spark by generating pipe between spark plug positive and negative polarities, lights the mixed gas of propellant. Described delay circuit (7) is arranged on the control end of oscillating circuit (4), when gas oxygen/kerosene rocket engine is lighted, after steady operation, cuts off oscillating circuit, stops getting angry. The present invention has obtained the reliable ignition under severe working environment, for small-scale structure, repeatedly start, the low-thrust rocket work of response and stable state/pulse double working modes provides guarantee fast.
Description
Technical field
The present invention relates to a kind of electric ignition device of spacecraft rocket engine.
Background technology
China's active service double elements low-thrust rocket is poisonous hypergolic propellants engine, and these havePoison hypergolic propellants in production, storage, transport and use procedure to natural environment and health structureBecome great harm. For conservation of nature environmental and human health impacts, carry out gas oxygen/kerosene propellant groupThe nontoxic low-thrust rocket development of closing. Because gas oxygen/kerosene propellant is combined as the two groups of nonflammableUnit's propellant, need to give engine configurations igniter. Conventionally the point that nonflammable rocket engine usesFire device has nature liquid dot firearm, solid gunpowder igniter and torch igniter etc., but gas oxygen/Kerosene low-thrust rocket is small-scale structure and need repeatedly starts, respond fast and stable state/pulse pairMode of operation, above-mentioned igniter is all inapplicable. Therefore, need to develop spacecraft rocket engine electricity pointFirearm.
Current domestic technical descriptioon or the report with this spacecraft rocket engine electric igniter that do not have to findRoad, also not yet collects both at home and abroad similarly data.
Summary of the invention
In order to solve the deficiency of existing rocket motor ignition technology, the object of the present invention is to provideA kind of spacecraft rocket engine electric igniter. Utilize the present invention, make igniter structure miniaturization, energyOn thrustor head space, install; Electric igniter is lightweight, can adapt to space product counterweightThe careless requirement at quarter of amount; Supply voltage 27VDC, meets supply voltage on carrier rocket arrow; Spark plug is putElectricity frequency is brought up to 200SPS, i.e. spark of 5ms meets appearance control rocket engine multiple-pulse/littlePulsed operation requirement; Delay circuit is set, lights and enter after steady operation state when engine, electric pointFirearm can automatically implement to cut out without shutdown signal in the situation that, plays the effect of protection electric igniter;Electric igniter and Propellant Control valve are controlled simultaneously, have simplified and the interface relationship of control system, favourableApplication in electric igniter on rocket engine.
In order to reach foregoing invention object, the technical scheme that the present invention adopts for its technical problem of solutionBe to provide a kind of spacecraft rocket engine electric igniter, this device comprises: spark plug shields by high pressureCovering cable is connected with driving source; Described driving source comprises that an oscillating circuit turns the DC voltage of inputChange high-frequency ac voltage into; A step-up transformer boosts to this high-frequency ac voltage; One wholeStream, energy storage, discharge circuit convert high-frequency ac voltage to high direct voltage and put on the positive and negative of spark plugThe two poles of the earth; Described rectification, energy storage, discharge circuit comprises commutation diode, energy storage capacitor and generating pipe,Connect with spark plug by high-tension shielding cable; Described spark plug is semiconductor-type spark plug; DescribedThe control end of oscillating circuit is provided with a delay circuit; The high-frequency ac that described step-up transformer is secondaryVoltage obtains 2000V DC voltage by described rectifies, and described energy storage capacitor is enteredRow charging, when charging reaches certain voltage value, energy storage capacitor is managed by generating and high-tension shielding cableTo described spark plug electric discharge, between spark plug positive and negative polarities, produce spark, light the mixed gas of propellant;Described delay circuit is lighted at gas oxygen/kerosene rocket engine, enters duty, time delay 120msRear cut-out oscillating circuit, stops spark plug and gets angry. Spark plug and driving source can be split type or oneFormula, driving source also can connect one or two spark plug, the invention process by high-tension shielding cableWhat example adopted is split type, and a spark plug connects the structure of a driving source. Electric igniter spark plugCan also be semiconductor-type or high voltage type, what the embodiment of the present invention adopted be semiconductor-type spark plug. InstituteState the metal material that spark-plug shell and electrode adopt high temperature resistant resistance to chemical attack, in spark-plug body, adoptInsulate with high alumina ceramic; By semiconductive ceramic creeping discharge, produce the mixed gas of spark ignition propellant,Spark plug discharge frequency reaches 200SPS. Be provided with a delay circuit at the control end of oscillating circuit,This circuit adopts 555 timers, disconnects, thereby reach a little for the time delay of controlling oscillating circuitThe control of firearm working time. The delay time arranging is auto-breaking after 120ms, and spark plug stopsGet angry.
The present invention's spacecraft rocket engine electric igniter is owing to taking miniaturization structure, DCElectricity, reduce driving source secondary voltage, improve spark plug discharge frequency, set up delay circuit and adopt halfThe technical schemes such as conductor type spark plug, have realized application on low-thrust rocket, have improvedIgniting security, extended the service life of device, to get angry end pollution insensitive, obtainedReliable ignition under severe working environment is the stable state/pulse of spacecraft low-thrust rocketThe double working modes beneficial effect such as give security.
Brief description of the drawings
Fig. 1 is the composition schematic diagram of spacecraft rocket engine electric igniter of the present invention.
Fig. 2 is the circuit structure block diagram of the driving source of electric igniter.
Detailed description of the invention
Below in conjunction with drawings and Examples, the present invention is further detailed explanation.
As shown in Figure 1, spark plug 1 is connected with driving source 3 by high-tension shielding cable 2. Electric ignitionThe driving source of device can connect one or two spark plug by high-tension shielding cable. The embodiment of the present inventionWhat adopt is semiconductor-type spark plug, and spark plug 1 shell and electrode adopt high temperature resistant resistance to chemical attackMetal material, adopts high alumina ceramic insulation, heat resistance and good airproof performance in spark plug 1 housing; AdoptSemiconductive ceramic creeping discharge technology produces spark, lights the mixed gas of propellant; Due to semiconductor-type sparkPlug has strong arcing ability, and impulsive force is large, can blow down foreign material, and reliable ignition is spacecraft low thrustStable state/pulse double working modes of rocket engine provides useful guarantee. Described spark plug 1 meetsMiniaturization structure designing requirement, the ignition end diameter of its spark plug 1 is Φ 6mm, length is 35mm,Driving source 3 weight < 0.5kg can install on thrustor head space. High-tension shielding electricityCable 2 is for connecting spark plug 1 and driving source 3, because the output of driving source produces 2000V direct currentHigh pressure adopts high-tension shielding cable 2 can make electric igniter have a good electricity high voltage bearing simultaneouslyMagnetic compatibility and antijamming capability. Driving source 3 adopts capacitor-discharge formula, driving source lightweight, energyAdapt to the careless quarter requirement of space product to weight.
Fig. 2 is the circuit structure block diagram of the driving source of electric igniter. As shown in Figure 2, space flight of the present inventionThe driving source 3 of device rocket engine electric igniter by oscillating circuit 4, step-up transformer 5, rectification,Energy storage, discharge circuit 6, delay circuit 7 form. Driving source 3 is directly powered by+27VDC, symbolClose the supply voltage on carrier rocket arrow. Oscillating circuit 4 converts input+27V DC voltage to heightFrequently alternating voltage, step-up transformer 5 is input to rectification, energy storage, puts after this alternating voltage is boostedElectricity circuit 6. Above-mentioned rectification, energy storage, discharge circuit 6 comprise commutation diode, energy storage capacitor and send outFulgurite, connects with spark plug 1 by high-tension shielding cable 2. The alternating voltage that step-up transformer is secondaryObtain 2000V high direct voltage by diode rectification, above-mentioned energy storage capacitor is charged, work as electric capacityMagnitude of voltage rise to when setting value, energy storage capacitor discharges to spark plug 1 by generating pipe,Between semiconductor-type spark plug positive and negative polarities, produce the voltage of 2000V, semiconductor creeping discharge producesSpark. Spark frequency reaches 200SPS, and spark of 5ms, meets appearance control rocket engine manyThe job requirement of pulse/small-pulse effect. At the control end of oscillating circuit 4, be also provided with a delay circuit7, this circuit adopts 555 timers, and delay time is 120ms, for controlling oscillating circuit 4Time delay disconnect, thereby reach the control to the electric igniter working time; Light and enter surely when engineAfter state duty, electric igniter can automatically implement to cut out without shutdown signal in the situation that, plays guarantorProtect the effect of electric igniter. The valve of above-mentioned igniter input direct voltage and engine uses synchronizes controlSignal processed, is powered by same power supply+27VDC, and electric igniter and Propellant Control valve are controlled simultaneously,Simplify and the interface relationship of control system, be conducive to the application of electric igniter on rocket engine.
Claims (10)
1. a spacecraft rocket engine electric igniter, is characterized in that, this device comprises:
Spark plug (1) is connected with driving source (3) by high-tension shielding cable (2); Described driving source(3) comprise that an oscillating circuit (4) converts the DC voltage of input to high-frequency ac voltage; OneIndividual step-up transformer (5) boosts to this high-frequency ac voltage; Rectification, energy storage, electric discharge electricityRoad (6) converts high-frequency ac voltage high direct voltage to and puts on the positive and negative polarities of spark plug (1);Described rectification, energy storage, discharge circuit (6) comprise commutation diode, energy storage capacitor and generating pipe,Connect by the same spark plug of high-tension shielding cable (2) (1); The control end of described oscillating circuit (4)Be provided with a delay circuit (7); The secondary high-frequency ac voltage of described step-up transformer (5) passes throughDescribed rectifies obtains 2000V DC voltage, and described energy storage capacitor is charged, and fillsWhen electricity reaches certain voltage value, energy storage capacitor manage by generating and high-tension shielding cable (2) to describedSpark plug (1) electric discharge produces spark between spark plug (1) positive and negative polarities, lights the mixed gas of propellant;Described delay circuit (7) cuts off oscillating circuit after gas oxygen/kerosene rocket engine is lighted steady operation(4), stopping spark plug (1) gets angry.
2. electric igniter according to claim 1, is characterized in that: described electric igniter also canTo be split type, do not have high-tension shielding cable to connect, spark plug connects the structure of driving source.
3. electric igniter according to claim 1, is characterized in that: described spark plug (1)For the semiconductor-type spark plug of high alumina ceramic insulation, spark-plug shell and electrode are high temperature resistant chemically-resistant corruptionThe metal material of erosion, adopts high alumina ceramic insulation in housing.
4. according to the electric igniter described in claim 1 or 3, it is characterized in that: spark plug (1)Ignition end diameter be Φ 6mm, length is 35mm, driving source (3) weight < 0.5kg.
5. electric igniter according to claim 1, is characterized in that: described electric igniterThe valve of input direct voltage and engine uses synchronous control signal, is powered by same power supply+27VDC.
6. electric igniter according to claim 1, is characterized in that: the electric discharge of electric igniter frequentlyRate is 200SPS, i.e. spark of 5ms.
7. electric igniter according to claim 1, is characterized in that: described delay circuit (7)Be 555 timers, be 120ms time delay turn-off time of setting.
8. electric igniter according to claim 1, is characterized in that: described spark plug (1)Adopt semiconductor-type spark plug.
9. electric igniter according to claim 1, is characterized in that: described spark plug (1)Adopt high voltage type spark plug.
10. electric igniter according to claim 1, is characterized in that: described electric igniterDriving source (3) can connect two spark plugs (1) by two high-tension shielding cables (2) simultaneously.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200610120358.3A CN106134428B (en) | 2006-10-19 | 2006-10-19 | Spacecraft rocket engine electric igniter |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN200610120358.3A CN106134428B (en) | 2006-10-19 | 2006-10-19 | Spacecraft rocket engine electric igniter |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN106134428B true CN106134428B (en) | 2011-06-01 |
Family
ID=57251139
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN200610120358.3A Expired - Fee Related CN106134428B (en) | 2006-10-19 | 2006-10-19 | Spacecraft rocket engine electric igniter |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN106134428B (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107939549A (en) * | 2017-11-08 | 2018-04-20 | 航宇救生装备有限公司 | A kind of miniature multistage pulses thrust solid propellant rocket |
| CN112304159A (en) * | 2020-10-29 | 2021-02-02 | 上海空间推进研究所 | Integrated gas supply device |
-
2006
- 2006-10-19 CN CN200610120358.3A patent/CN106134428B/en not_active Expired - Fee Related
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107939549A (en) * | 2017-11-08 | 2018-04-20 | 航宇救生装备有限公司 | A kind of miniature multistage pulses thrust solid propellant rocket |
| CN107939549B (en) * | 2017-11-08 | 2019-10-15 | 航宇救生装备有限公司 | A kind of miniature multistage pulses thrust solid propellant rocket |
| CN112304159A (en) * | 2020-10-29 | 2021-02-02 | 上海空间推进研究所 | Integrated gas supply device |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN104595087B (en) | Pass through the igniter of the fuel air mixture in corona discharge ignition engine | |
| EP1788235B1 (en) | Plasma jet spark plug and ignition system | |
| CN106640568B (en) | A kind of dipolar solid ablative-type protective coating plasma accelerator | |
| CA2584628C (en) | Triggered pulsed ignition system and method | |
| CN101463764B (en) | High-energy plasma igniter of gas turbine | |
| CN109404198A (en) | A kind of plasma generator | |
| CN103423061A (en) | Extra high-energy spark discharge system | |
| CN106134428B (en) | Spacecraft rocket engine electric igniter | |
| CN210122944U (en) | Ignition device with adjustable spark plug gap | |
| CN203464329U (en) | Boiler ignition device | |
| EP1214519A1 (en) | Long-life traveling spark ignitor and associated firing circuitry | |
| CN112963249A (en) | High-energy igniter for aviation | |
| CN102563693A (en) | High voltage insulation diesel spray ignition nozzle | |
| CN204100290U (en) | A kind of high-energy electronic igniter circuit | |
| CN204513450U (en) | A kind of firing circuit of ignition | |
| WO2001028053A1 (en) | Ignition system for an internal combustion engine | |
| CN104214797A (en) | High-energy electronic igniter circuit | |
| CN101844277A (en) | Discharging circuit for welding discharger | |
| EP0001354A1 (en) | Ignition system | |
| CN103821656A (en) | Multi-mode discharging ignition system for gas engine | |
| CN219431941U (en) | High-energy igniter for gas turbine | |
| CN103225583A (en) | Ignition device for gas engine | |
| CN202157890U (en) | Ignition device of motorcycle | |
| WO2004028207A3 (en) | Igniter curcuit with an air gap | |
| RU74667U1 (en) | COMBINED IGNITION SYSTEM |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| GR03 | Grant of secret patent right | ||
| DC01 | Secret patent status has been lifted | ||
| DCSP | Declassification of secret patent | ||
| CF01 | Termination of patent right due to non-payment of annual fee | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20110601 Termination date: 20201019 |