超级军网探索长征五号5米芯级模块火箭神秘的历史定位及其背后的 ...
来源：百度文库 编辑：超级军网 时间：2024/04/28 01:32:31
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扯了半天，终于回归楼主极度吹嘘的360吨上了;P
Manufacturer Name: NK-33. Government Designation: 11D111. Designer: Kuznetsov. Developed in: 1970-74. Application: N-1F stage 1 (block A). Propellants: Lox/Kerosene. Thrust(vac): 1,638.000 kN (368,237 lbf). Thrust(sl): 1,510.200 kN (339,506 lbf). Isp: 331 sec. Isp (sea level): 297 sec. Burn time: 600 sec. Mass Engine: 1,222 kg (2,694 lb). Diameter: 1.50 m (4.90 ft). Length: 3.71 m (12.17 ft). Chambers: 1. Chamber Pressure: 145.70 bar. Area Ratio: 27.00. Oxidizer to Fuel Ratio: 2.80. Thrust to Weight Ratio: 136.66. Country: Russia. Status: Design 1975.
After initial failures of N1 modified engines with multiple ignition capability and increased life were developed. NK-33 was tested at thrust of up to 204 t, variations of components ratio up to 20% and total burning time up to 1200 sec. Mothballed, never flown until selected by Kistler for their commercial launch vehicle in 1996.
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NK-33 used on Rocket Stages
Used on stage: Kistler Stage 1. on launch vehicle: Kistler K-1.
Used on stage: N1F Block A. on launch vehicle: N1F.
Used on stage: N1F Block A. on launch vehicle: N1F Sr.
Used on stage: N1F Block A. on launch vehicle: N1F-L3M.
Used on stage: UR-500MK-1. on launch vehicle: UR-500MK.
这是苏联当年的为N1火箭开发的NK33液氧煤油发动机,后来也最终白白放弃了.比较可惜.
After initial failures of N1 modified engines with multiple ignition capability and increased life were developed. NK-33 was tested at thrust of up to 204 t, variations of components ratio up to 20% and total burning time up to 1200 sec. Mothballed, never flown until selected by Kistler for their commercial launch vehicle in 1996.
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NK-33 used on Rocket Stages
Used on stage: Kistler Stage 1. on launch vehicle: Kistler K-1.
Used on stage: N1F Block A. on launch vehicle: N1F.
Used on stage: N1F Block A. on launch vehicle: N1F Sr.
Used on stage: N1F Block A. on launch vehicle: N1F-L3M.
Used on stage: UR-500MK-1. on launch vehicle: UR-500MK.
这是苏联当年的为N1火箭开发的NK33液氧煤油发动机,后来也最终白白放弃了.比较可惜.
虽然最后又是个无敌360
但是文中某些观点还是有些道理.:b
但是文中某些观点还是有些道理.:b
Orbital launch vehicle. Year: 1962. Family: R-36. Country: Ukraine. Status: Study 1961. Article Number: 8K68. Manufacturer's Designation: R-56. Complex: 8K68.
The R-56 was Yangel's ultimate superbooster design. Trade studies begun in 1962 resulted in a conventional tandem stage design capable of being transported on the Soviet canal system from the factory to the launch site, while still placing 40 metric tons into low earth orbit. However various Soviet government factions favored the much larger (and less practical) Korolev N1 or Chelomei UR-700 designs. Yangel made one last attempt to convince the government to sponsor a common approach to the lunar program, with different design bureaus concentrating on just one part of the mission, as the American's were doing. But his practical solutions obtained no traction, and further work on the R-56 was abandoned.
After drawing back from the 'cluster of R-16's' approach of the SK-100, Yangel conducted some trade studies to determine the optimum design for his bureau's first 憇uper rocket? The booster was to be capable of serving as a first-strike military global rocket or as a heavy launch vehicle, placing 40 metric ton payloads into a 200 km polar orbit. The selected monoblock design could be transported on the Soviet internal canal system from the factory to the launch site.
The R-56 would have been 67.8 m long and consisted of three stages, the first two with a basic diameter of 6.5 m. The first stage had a flared 8.2 m diameter base to accommodate the 16 RD-253 engines.
Yangel conducted some trade studies to determine the optimum design for his bureau's first 憇uper rocket? The booster was to be capable of launching a range of missions:
First-strike military global rocket: 35 metric tons of independently maneuverable orbital nuclear warheads over a 16,000 km range
Heavy launch vehicle for military space stations or weapons platforms: 40 metric ton payload into a 200 km polar orbit
Global communications satellites in geosynchronous or inclined pseudo-synchronous (figure 8) orbits
Unmanned lunar surface monitoring stations
Manned circumlunar and lunar orbital mapping missions: 12 metric tons into lunar orbit
Through use of docking in low earth orbit, assembly of spacecraft for manned lunar expeditions
High priority payloads for manned lunar expeditions
Large unmanned interplanetary probes: payloads of 6 to 8 metric tons on trajectories to Mars or Venus
To achieve these design objectives, three design approaches were studied in detail:
A polyblock design limited to rail transport restrictions (4 x 3.8 m diameter stages clustered together)
A polyblock design using existing R-36 ICBM tooling (7 x 3.0 m diameter stages clustered together, similar to the SK-100)
A monoblock design that could be transported by waterway (6.5 m diameter)
Yangel reached a similar conclusion to that of Korolev in the draft project for the N1. The two stage monoblock design was clearly superior to the polyblock versions. Among its advantages:
Less labor required for manufacture at the plant and integration at the launch site
Design solutions already used on R-16/R-36 ICBM's could be scaled up
Vehicle dynamics models developed for R-16/R-36 could be used
Less development work due to less complex vehicle dynamics
Lower launch complex cost due to reduced number of fuelling and interface points
Larger diameter core better suited to carrying low-density liquid hydrogen / nuclear thermal engine upper stages for manned interplanetary spacecraft.
The selected monoblock vehicle was 67.8 m long and consisted of three stages, the first two with a basic diameter of 6.5 m. The first stage had a flared 8.2 m diameter base to accommodate the 16 engines. These were developed by Glushko's OKB-456, and each produced 148 metric tons of thrust at lift-off. The thrust and technical characteristics indicate they would have been identical to or closely derived from the RD-253 engine developed for Chelomei's UR-500 rocket. Twelve of the main engines were fixed while four were gimbaled in pitch to provide launch vehicle steering. Separate propellant tanks were arranged with the oxidizer forward tank forward, the fuel tank aft. It is said that the launch vehicle was equipped with a recovery system; this probably applied only to the first stage, which the figures indicate has an unusually high empty mass, perhaps including a means of recovering the stage or engine section for reuse.
The second stage had a common bulkhead between the oxidizer and propellant tanks. It was equipped with a single fixed main engine of 172 metric tons thrust. The technical characteristics indicate that it was identical with the RD-254 high-expansion ration derivative of the RD-253 also developed for Chelomei's UR-700. Four vernier engines, with a total thrust of 50,000 kgf, provided ullage force for main engine ignition, steering during main engine burn, and precise velocity correction after main engine cut-off.
There were two versions of the third stage: a single start variant for placement of large payloads into low earth orbit, and a multiple start version for taking smaller payloads to high earth orbit or earth escape trajectories. It was equipped with a main engine of 50,000 kgf and four vernier engines totaling 5,500 kgf. The characteristics indicate it was probably based on Kosberg's RD-0213 engine developed for the third stage of the UR-500K Proton launch vehicle. The stage, while having nearly the same diameter as the Proton third stage, differed in design, with a common bulkhead between the oxidizer and propellant tanks.
A fourth stage was designed to allow insertion of payloads into geosynchronous or lunar orbit. The propellants used and engine characteristics are identical to those of the experimental RD-280 engine developed by Glushko in the mid-1960's. The stage uniquely used the Aerozine-50 fuel developed in the United States for the Titan 2 ICBM. The engine was capable of four restarts.
The assembled vehicle would be moved by road from the Yangel factory in Dnepropetrovsk to the mouth of the Surya River. From there it could be moved on the Soviet Union's elaborate inland waterway system to any of the existing launch sites - Kapustin Yar on the Volga, Baikonur on the Syr Darya, or Plesetsk on the Yemtsa. Kapustin Yar, the closest to the factory, was the preferred location. At the launch site the vehicle would be moved to an austere launch pad a short distance form the river. After being put upright it would be enclosed by a service tower providing an environmentally protected environment for final rocket and payload preparation.
Growth versions of the R-56 would return to the cluster principle in order to boost larger payloads, such as those required for manned lunar expeditions. The 685,000 kgf RD-270 engine was also to have been used in later versions of the R-56. This would have reduced the number of engines in the first stage from 16 to 4.
Yangel's KB Yuzhnoye was the prime contractor for the launch vehicle, with support from the Soviet Academy of Sciences, NII-88, NITI-40, GSPI, and the Ministry of Defense. By 1965 Yangel had decided that the bitter fight between Chelomei and Korolev over control of manned programs was damaging the Soviet space effort. In any case he could see that the size of the projects had reached such a scale that it was impossible for one design bureau to handle all of the required elements. He proposed a collaborative effort: Yangel would design and build the launch vehicle; Korolev the manned spacecraft; and Chelomei the unmanned spacecraft.
However this was not to be. The leadership was loath to change course with funds already invested in development of boosters and spacecraft by Chelomei and Korolev. The other Chief Designers objected that use of the R-56 for a manned lunar landing would require two R-56 launches in the place of one UR-700 or N1 launch. This would mean use of untried earth orbit rendezvous techniques to assemble the spacecraft in earth orbit. Development of the R-56 was not authorized, and for once in his career Yangel gave up the fight.
As a practical matter it was not possible for one bureau to handle the moon landing project. Although Glushko and Chelomei refused to co-operate with Korolev on the effort, most other rocket design bureaus were involved. Yangel found himself charged with development of the LK lunar lander that would actually land a cosmonaut on the moon.
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Many thanks to Asif Siddiqi and Marcus Lindroos for providing source materials for this article.
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Manufacturer: Yuzhnoye. LEO Payload: 40,000 kg (88,000 lb). to: 200 km Orbit. at: 90.00 degrees. Payload: 6,000 kg (13,200 lb). to a: geosynchronous orbit trajectory. Liftoff Thrust: 23,220.000 kN (5,220,060 lbf). Total Mass: 1,421,000 kg (3,132,000 lb). Core Diameter: 6.50 m (21.30 ft). Total Length: 67.80 m (222.40 ft).
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Stage Data - R-56
Stage Number: 1. 1 x Stage: R-56 Block A. Gross Mass: 1,162,000 kg (2,561,000 lb). Empty Mass: 162,000 kg (357,000 lb). Thrust (vac): 1,608.000 kN (361,492 lbf). Isp: 316 sec. Burn time: 120 sec. Isp(sl): 285 sec. Diameter: 6.50 m (21.30 ft). Span: 8.20 m (26.90 ft). Length: 36.00 m (118.00 ft). Propellants: N2O4/UDMH. No Engines: 16. Engine: RD-253-11D48. Status: Study 1961.
Stage Number: 2. 1 x Stage: R-56 Block B. Gross Mass: 213,000 kg (469,000 lb). Empty Mass: 13,000 kg (28,000 lb). Thrust (vac): 1,980.000 kN (445,120 lbf). Isp: 325 sec. Burn time: 370 sec. Diameter: 6.50 m (21.30 ft). Span: 6.50 m (21.30 ft). Length: 12.00 m (39.00 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-254. Status: Study 1961.
Stage Number: 3. 1 x Stage: R-56 Block O. Gross Mass: 32,000 kg (70,000 lb). Empty Mass: 2,000 kg (4,400 lb). Thrust (vac): 544.200 kN (122,341 lbf). Isp: 327 sec. Burn time: 190 sec. Diameter: 4.00 m (13.10 ft). Span: 4.00 m (13.10 ft). Length: 7.00 m (22.90 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-0213. Other designations: Orbital Block. Status: Study 1961. Empty mass estimated (rounded figures given in source material indicates impossible 31 tonnes gross with 30 tonnes propellant).
Stage Number: 4. 1 x Stage: R-56 Block K. Gross Mass: 9,400 kg (20,700 lb). Empty Mass: 700 kg (1,540 lb). Thrust (vac): 117.000 kN (26,302 lbf). Isp: 350 sec. Burn time: 350 sec. Diameter: 4.00 m (13.10 ft). Span: 4.00 m (13.10 ft). Length: 5.00 m (16.40 ft). Propellants: N2O4/Aerozine-50. No Engines: 1. Engine: RD-280. Other designations: Space Block. Status: Study 1961. Empty mass estimated (rounded figures given in source material indicates impossible 9 tonnes gross with 8.7 tonnes propellant).
这是苏联杨格列当年的R56,起飞推力2200吨左右,LEO运载能力40吨,这也是一种在载人环月飞行任务中大有作为的火箭.
The R-56 was Yangel's ultimate superbooster design. Trade studies begun in 1962 resulted in a conventional tandem stage design capable of being transported on the Soviet canal system from the factory to the launch site, while still placing 40 metric tons into low earth orbit. However various Soviet government factions favored the much larger (and less practical) Korolev N1 or Chelomei UR-700 designs. Yangel made one last attempt to convince the government to sponsor a common approach to the lunar program, with different design bureaus concentrating on just one part of the mission, as the American's were doing. But his practical solutions obtained no traction, and further work on the R-56 was abandoned.
After drawing back from the 'cluster of R-16's' approach of the SK-100, Yangel conducted some trade studies to determine the optimum design for his bureau's first 憇uper rocket? The booster was to be capable of serving as a first-strike military global rocket or as a heavy launch vehicle, placing 40 metric ton payloads into a 200 km polar orbit. The selected monoblock design could be transported on the Soviet internal canal system from the factory to the launch site.
The R-56 would have been 67.8 m long and consisted of three stages, the first two with a basic diameter of 6.5 m. The first stage had a flared 8.2 m diameter base to accommodate the 16 RD-253 engines.
Yangel conducted some trade studies to determine the optimum design for his bureau's first 憇uper rocket? The booster was to be capable of launching a range of missions:
First-strike military global rocket: 35 metric tons of independently maneuverable orbital nuclear warheads over a 16,000 km range
Heavy launch vehicle for military space stations or weapons platforms: 40 metric ton payload into a 200 km polar orbit
Global communications satellites in geosynchronous or inclined pseudo-synchronous (figure 8) orbits
Unmanned lunar surface monitoring stations
Manned circumlunar and lunar orbital mapping missions: 12 metric tons into lunar orbit
Through use of docking in low earth orbit, assembly of spacecraft for manned lunar expeditions
High priority payloads for manned lunar expeditions
Large unmanned interplanetary probes: payloads of 6 to 8 metric tons on trajectories to Mars or Venus
To achieve these design objectives, three design approaches were studied in detail:
A polyblock design limited to rail transport restrictions (4 x 3.8 m diameter stages clustered together)
A polyblock design using existing R-36 ICBM tooling (7 x 3.0 m diameter stages clustered together, similar to the SK-100)
A monoblock design that could be transported by waterway (6.5 m diameter)
Yangel reached a similar conclusion to that of Korolev in the draft project for the N1. The two stage monoblock design was clearly superior to the polyblock versions. Among its advantages:
Less labor required for manufacture at the plant and integration at the launch site
Design solutions already used on R-16/R-36 ICBM's could be scaled up
Vehicle dynamics models developed for R-16/R-36 could be used
Less development work due to less complex vehicle dynamics
Lower launch complex cost due to reduced number of fuelling and interface points
Larger diameter core better suited to carrying low-density liquid hydrogen / nuclear thermal engine upper stages for manned interplanetary spacecraft.
The selected monoblock vehicle was 67.8 m long and consisted of three stages, the first two with a basic diameter of 6.5 m. The first stage had a flared 8.2 m diameter base to accommodate the 16 engines. These were developed by Glushko's OKB-456, and each produced 148 metric tons of thrust at lift-off. The thrust and technical characteristics indicate they would have been identical to or closely derived from the RD-253 engine developed for Chelomei's UR-500 rocket. Twelve of the main engines were fixed while four were gimbaled in pitch to provide launch vehicle steering. Separate propellant tanks were arranged with the oxidizer forward tank forward, the fuel tank aft. It is said that the launch vehicle was equipped with a recovery system; this probably applied only to the first stage, which the figures indicate has an unusually high empty mass, perhaps including a means of recovering the stage or engine section for reuse.
The second stage had a common bulkhead between the oxidizer and propellant tanks. It was equipped with a single fixed main engine of 172 metric tons thrust. The technical characteristics indicate that it was identical with the RD-254 high-expansion ration derivative of the RD-253 also developed for Chelomei's UR-700. Four vernier engines, with a total thrust of 50,000 kgf, provided ullage force for main engine ignition, steering during main engine burn, and precise velocity correction after main engine cut-off.
There were two versions of the third stage: a single start variant for placement of large payloads into low earth orbit, and a multiple start version for taking smaller payloads to high earth orbit or earth escape trajectories. It was equipped with a main engine of 50,000 kgf and four vernier engines totaling 5,500 kgf. The characteristics indicate it was probably based on Kosberg's RD-0213 engine developed for the third stage of the UR-500K Proton launch vehicle. The stage, while having nearly the same diameter as the Proton third stage, differed in design, with a common bulkhead between the oxidizer and propellant tanks.
A fourth stage was designed to allow insertion of payloads into geosynchronous or lunar orbit. The propellants used and engine characteristics are identical to those of the experimental RD-280 engine developed by Glushko in the mid-1960's. The stage uniquely used the Aerozine-50 fuel developed in the United States for the Titan 2 ICBM. The engine was capable of four restarts.
The assembled vehicle would be moved by road from the Yangel factory in Dnepropetrovsk to the mouth of the Surya River. From there it could be moved on the Soviet Union's elaborate inland waterway system to any of the existing launch sites - Kapustin Yar on the Volga, Baikonur on the Syr Darya, or Plesetsk on the Yemtsa. Kapustin Yar, the closest to the factory, was the preferred location. At the launch site the vehicle would be moved to an austere launch pad a short distance form the river. After being put upright it would be enclosed by a service tower providing an environmentally protected environment for final rocket and payload preparation.
Growth versions of the R-56 would return to the cluster principle in order to boost larger payloads, such as those required for manned lunar expeditions. The 685,000 kgf RD-270 engine was also to have been used in later versions of the R-56. This would have reduced the number of engines in the first stage from 16 to 4.
Yangel's KB Yuzhnoye was the prime contractor for the launch vehicle, with support from the Soviet Academy of Sciences, NII-88, NITI-40, GSPI, and the Ministry of Defense. By 1965 Yangel had decided that the bitter fight between Chelomei and Korolev over control of manned programs was damaging the Soviet space effort. In any case he could see that the size of the projects had reached such a scale that it was impossible for one design bureau to handle all of the required elements. He proposed a collaborative effort: Yangel would design and build the launch vehicle; Korolev the manned spacecraft; and Chelomei the unmanned spacecraft.
However this was not to be. The leadership was loath to change course with funds already invested in development of boosters and spacecraft by Chelomei and Korolev. The other Chief Designers objected that use of the R-56 for a manned lunar landing would require two R-56 launches in the place of one UR-700 or N1 launch. This would mean use of untried earth orbit rendezvous techniques to assemble the spacecraft in earth orbit. Development of the R-56 was not authorized, and for once in his career Yangel gave up the fight.
As a practical matter it was not possible for one bureau to handle the moon landing project. Although Glushko and Chelomei refused to co-operate with Korolev on the effort, most other rocket design bureaus were involved. Yangel found himself charged with development of the LK lunar lander that would actually land a cosmonaut on the moon.
--------------------------------------------------------------------------------
Many thanks to Asif Siddiqi and Marcus Lindroos for providing source materials for this article.
--------------------------------------------------------------------------------
Manufacturer: Yuzhnoye. LEO Payload: 40,000 kg (88,000 lb). to: 200 km Orbit. at: 90.00 degrees. Payload: 6,000 kg (13,200 lb). to a: geosynchronous orbit trajectory. Liftoff Thrust: 23,220.000 kN (5,220,060 lbf). Total Mass: 1,421,000 kg (3,132,000 lb). Core Diameter: 6.50 m (21.30 ft). Total Length: 67.80 m (222.40 ft).
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Stage Data - R-56
Stage Number: 1. 1 x Stage: R-56 Block A. Gross Mass: 1,162,000 kg (2,561,000 lb). Empty Mass: 162,000 kg (357,000 lb). Thrust (vac): 1,608.000 kN (361,492 lbf). Isp: 316 sec. Burn time: 120 sec. Isp(sl): 285 sec. Diameter: 6.50 m (21.30 ft). Span: 8.20 m (26.90 ft). Length: 36.00 m (118.00 ft). Propellants: N2O4/UDMH. No Engines: 16. Engine: RD-253-11D48. Status: Study 1961.
Stage Number: 2. 1 x Stage: R-56 Block B. Gross Mass: 213,000 kg (469,000 lb). Empty Mass: 13,000 kg (28,000 lb). Thrust (vac): 1,980.000 kN (445,120 lbf). Isp: 325 sec. Burn time: 370 sec. Diameter: 6.50 m (21.30 ft). Span: 6.50 m (21.30 ft). Length: 12.00 m (39.00 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-254. Status: Study 1961.
Stage Number: 3. 1 x Stage: R-56 Block O. Gross Mass: 32,000 kg (70,000 lb). Empty Mass: 2,000 kg (4,400 lb). Thrust (vac): 544.200 kN (122,341 lbf). Isp: 327 sec. Burn time: 190 sec. Diameter: 4.00 m (13.10 ft). Span: 4.00 m (13.10 ft). Length: 7.00 m (22.90 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-0213. Other designations: Orbital Block. Status: Study 1961. Empty mass estimated (rounded figures given in source material indicates impossible 31 tonnes gross with 30 tonnes propellant).
Stage Number: 4. 1 x Stage: R-56 Block K. Gross Mass: 9,400 kg (20,700 lb). Empty Mass: 700 kg (1,540 lb). Thrust (vac): 117.000 kN (26,302 lbf). Isp: 350 sec. Burn time: 350 sec. Diameter: 4.00 m (13.10 ft). Span: 4.00 m (13.10 ft). Length: 5.00 m (16.40 ft). Propellants: N2O4/Aerozine-50. No Engines: 1. Engine: RD-280. Other designations: Space Block. Status: Study 1961. Empty mass estimated (rounded figures given in source material indicates impossible 9 tonnes gross with 8.7 tonnes propellant).
这是苏联杨格列当年的R56,起飞推力2200吨左右,LEO运载能力40吨,这也是一种在载人环月飞行任务中大有作为的火箭.
Orbital launch vehicle. Year: 1967. Family: UR. Country: Russia. Status: Out of production. Other Designations: Proton / Block D. Library of Congress Designation: D-1e. Department of Defence Designation: SL-12. Article Number: 8K82K. Manufacturer's Designation: UR-500K.
This four stage version of the Proton was originally designed to send manned circumlunar spacecraft into translunar trajectory. Guidance to the Block D stage must be supplied by spacecraft. The design was proposed on 8 September 1965 by Korolev as an alternate to Chelomei's LK-1 circumlunar mission. It combined the Proton 8K82K booster for the LK-1 with the N1 lunar Block D stage to boost a stripped-down Soyuz 7K-L1 spacecraft around the moon. The Korolev design was selected, and first flight came on 10 March 1967. The crash lunar program led to a poor launch record. Following a protracted ten year test period, the booster finally reached a level of launch reliability comparable to that of other world launch vehicles.
Development of a three-stage version of the UR-500 was authorised in the decree of 3 August 1964. Decrees of 12 October and 11 November 1964 authorised development of the Almaz manned military space station and the manned circumlunar spacecraft LK-1 as payloads for the UR-500K. However at the same time Khrushchev was ousted from power. Chelomei lost his chief patron and his projects came under negative scrutiny by the new leadership.
On 8 September 1965 Korolev presented an several schemes for using Chelomei抯 UR-500 to fly around the moon. One alternate was a two-part spaceship, using the Proton with the upper stage Block D from Korolev抯 N1-L3 lunar project. This would launch Korolev抯 7K-L1 spacecraft (derived from the 7K-OK Soyuz) onto a translunar trajectory. This project received the name UR-500K-L1, and was adopted in place of Chelomei抯 LK-1 circumlunar project. It required construction of 18 UR-500K rockets, which, in a combination flight-test and government trials program, would send L1 spacecraft around the moon, at first unmanned, then manned.
By 4 October 1966 a dummy rocket was mounted at the launch site. The dummy was loaded with imitation propellants (kerosene as fuel and water/ethyl alcohol as oxidiser). The nitrogen tetroxide oxidiser had to be kept above -11 degrees C, and it was originally planned for a thermostatically-controlled electrical heating of the tank walls to achieve this. It was ultimately decided that the risk of explosion of such a system was too great, and the system was abandoned.
The first flight rocket (serial number 22701) began assembly on 21 November 1966, with mechanical assembly completed by 29 November. Electrical connections and tests were completed by 4 December 1966. Due to New Year抯 holidays work did not resume until 28 January 1967. By 28 February the fully assembled booster / spacecraft unit was completed in the MIK, including the 7K-L1P boilerplate spacecraft. The launch tower was added on 2 March 1967 and the system was declared ready for launch. A serious potential problem during preparations was the discovery that fuel gases could lead to pump cavitation at the turbine exits. Tests on the ground showed that the problem was not the fuel itself, but in the monitoring equipment.
Although the first launch of the UR-500K-L1 on 10 March was successful, the record for the balance of the manned circumlunar project was dismal. Of the remaining 11 launches of the project, only that of Zond-7 was recognised as fully successful. In 60% of the failures the fault was in the launch vehicle; in 20% the Block D; and in 20% the spacecraft. Therefore the probability of successfully carrying out the objective of the project - safely flying a cosmonaut around the moon and returning him to earth - was only 9%.
Remarkably, due to continuing failures, the 8K82K did not satisfactorily complete its state trials until its 61st launch (Salyut 6 / serial number 29501 / 29 September 1977). Thereafter it reached a level of launch reliability comparable to that of other world launch vehicles.
Manufacturer: Chelomei. Launches: 39. Failures: 15. Success Rate: 61.54%. First Launch Date: 1967-03-10. Last Launch Date: 1975-10-16. Launch data is: complete. Payload: 5,390 kg (11,880 lb). to a: translunar trajectory. Associated Spacecraft: Luna Ye-8, Luna Ye-8-5, Luna Ye-8-5M, Luna Ye-8-LS, Mars M-69, Mars M-71, Mars M-73, Soyuz 7K-L1 , Soyuz 7K-L1E, Soyuz 7K-L1P , Venera 4V-1, DLB Beacon Lander. Liftoff Thrust: 8,847.000 kN (1,988,884 lbf). Total Mass: 707,170 kg (1,559,040 lb). Core Diameter: 4.15 m (13.61 ft). Total Length: 57.00 m (187.00 ft). Launch Price $: 70.000 million. in: 1994 price dollars.
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Model: Proton-K/D. Family: UR. Country: Russia.
Production version.
LEO Payload: 5,000 kg (11,000 lb). Payload: 4,350 kg (9,590 lb). to a: geosynchronous transfer orbit trajectory. Apogee: 400,000 km (240,000 mi). Liftoff Thrust: 9,500.000 kN (2,135,600 lbf). Total Mass: 691,500 kg (1,524,400 lb). Core Diameter: 7.40 m (24.20 ft). Total Length: 55.40 m (181.70 ft).
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Model: UR-500K/Blok D. Family: UR. Country: Russia.
Prototype version.
LEO Payload: 5,000 kg (11,000 lb). Payload: 4,350 kg (9,590 lb). to a: geosynchronous transfer orbit trajectory. Apogee: 400,000 km (240,000 mi). Liftoff Thrust: 9,500.000 kN (2,135,600 lbf). Total Mass: 691,500 kg (1,524,400 lb). Core Diameter: 7.40 m (24.20 ft). Total Length: 55.40 m (181.70 ft).
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Stage Data - Proton 8K82K / 11S824
Stage Number: 1. 1 x Stage: Proton K-1. Gross Mass: 450,510 kg (993,200 lb). Empty Mass: 31,100 kg (68,500 lb). Thrust (vac): 10,470.158 kN (2,353,785 lbf). Isp: 316 sec. Burn time: 124 sec. Isp(sl): 267 sec. Diameter: 4.15 m (13.61 ft). Span: 7.40 m (24.20 ft). Length: 21.20 m (69.50 ft). Propellants: N2O4/UDMH. No Engines: 6. Engine: RD-253-11D48. Other designations: 8S810K.
Stage Number: 2. 1 x Stage: Proton K-2. Gross Mass: 167,828 kg (369,997 lb). Empty Mass: 11,715 kg (25,827 lb). Thrust (vac): 2,399.216 kN (539,365 lbf). Isp: 327 sec. Burn time: 206 sec. Isp(sl): 230 sec. Diameter: 4.15 m (13.61 ft). Span: 4.15 m (13.61 ft). Length: 14.00 m (45.00 ft). Propellants: N2O4/UDMH. No Engines: 4. Engine: RD-0210. Other designations: 8S811K.
Stage Number: 3. 1 x Stage: Proton K-3. Gross Mass: 50,747 kg (111,877 lb). Empty Mass: 4,185 kg (9,226 lb). Thrust (vac): 630.170 kN (141,668 lbf). Isp: 325 sec. Burn time: 238 sec. Isp(sl): 230 sec. Diameter: 4.15 m (13.61 ft). Span: 4.15 m (13.61 ft). Length: 6.50 m (21.30 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-0212.
Stage Number: 4. 1 x Stage: Proton 11S824. Gross Mass: 13,360 kg (29,450 lb). Empty Mass: 1,800 kg (3,900 lb). Thrust (vac): 83.300 kN (18,727 lbf). Isp: 346 sec. Burn time: 470 sec. Diameter: 3.70 m (12.10 ft). Span: 3.70 m (12.10 ft). Length: 5.50 m (18.00 ft). Propellants: Lox/Kerosene. No Engines: 1. Engine: RD-58. Other designations: 11S824; Block D; D-1-e. Block D, article number 11S824. Without guidance unit (navigation commands come from payload). Originally designed as N1-L3 lunar expedition launch vehicle lunar orbit insertion/lunar crasher stage. Adapted for use with Proton UR-500K as a fourth stage for manned circumlunar flight. Further used to launch large Lavochkin bureau unmanned lunar/planetary spacecraft. Flown from 1967 to 1975.
这是当年苏联准备用于载人环月飞行的质子火箭,但最终由于成功率太低,把握不大,又放弃了.这一放弃,苏联人就彻底与月球失之交臂了.
This four stage version of the Proton was originally designed to send manned circumlunar spacecraft into translunar trajectory. Guidance to the Block D stage must be supplied by spacecraft. The design was proposed on 8 September 1965 by Korolev as an alternate to Chelomei's LK-1 circumlunar mission. It combined the Proton 8K82K booster for the LK-1 with the N1 lunar Block D stage to boost a stripped-down Soyuz 7K-L1 spacecraft around the moon. The Korolev design was selected, and first flight came on 10 March 1967. The crash lunar program led to a poor launch record. Following a protracted ten year test period, the booster finally reached a level of launch reliability comparable to that of other world launch vehicles.
Development of a three-stage version of the UR-500 was authorised in the decree of 3 August 1964. Decrees of 12 October and 11 November 1964 authorised development of the Almaz manned military space station and the manned circumlunar spacecraft LK-1 as payloads for the UR-500K. However at the same time Khrushchev was ousted from power. Chelomei lost his chief patron and his projects came under negative scrutiny by the new leadership.
On 8 September 1965 Korolev presented an several schemes for using Chelomei抯 UR-500 to fly around the moon. One alternate was a two-part spaceship, using the Proton with the upper stage Block D from Korolev抯 N1-L3 lunar project. This would launch Korolev抯 7K-L1 spacecraft (derived from the 7K-OK Soyuz) onto a translunar trajectory. This project received the name UR-500K-L1, and was adopted in place of Chelomei抯 LK-1 circumlunar project. It required construction of 18 UR-500K rockets, which, in a combination flight-test and government trials program, would send L1 spacecraft around the moon, at first unmanned, then manned.
By 4 October 1966 a dummy rocket was mounted at the launch site. The dummy was loaded with imitation propellants (kerosene as fuel and water/ethyl alcohol as oxidiser). The nitrogen tetroxide oxidiser had to be kept above -11 degrees C, and it was originally planned for a thermostatically-controlled electrical heating of the tank walls to achieve this. It was ultimately decided that the risk of explosion of such a system was too great, and the system was abandoned.
The first flight rocket (serial number 22701) began assembly on 21 November 1966, with mechanical assembly completed by 29 November. Electrical connections and tests were completed by 4 December 1966. Due to New Year抯 holidays work did not resume until 28 January 1967. By 28 February the fully assembled booster / spacecraft unit was completed in the MIK, including the 7K-L1P boilerplate spacecraft. The launch tower was added on 2 March 1967 and the system was declared ready for launch. A serious potential problem during preparations was the discovery that fuel gases could lead to pump cavitation at the turbine exits. Tests on the ground showed that the problem was not the fuel itself, but in the monitoring equipment.
Although the first launch of the UR-500K-L1 on 10 March was successful, the record for the balance of the manned circumlunar project was dismal. Of the remaining 11 launches of the project, only that of Zond-7 was recognised as fully successful. In 60% of the failures the fault was in the launch vehicle; in 20% the Block D; and in 20% the spacecraft. Therefore the probability of successfully carrying out the objective of the project - safely flying a cosmonaut around the moon and returning him to earth - was only 9%.
Remarkably, due to continuing failures, the 8K82K did not satisfactorily complete its state trials until its 61st launch (Salyut 6 / serial number 29501 / 29 September 1977). Thereafter it reached a level of launch reliability comparable to that of other world launch vehicles.
Manufacturer: Chelomei. Launches: 39. Failures: 15. Success Rate: 61.54%. First Launch Date: 1967-03-10. Last Launch Date: 1975-10-16. Launch data is: complete. Payload: 5,390 kg (11,880 lb). to a: translunar trajectory. Associated Spacecraft: Luna Ye-8, Luna Ye-8-5, Luna Ye-8-5M, Luna Ye-8-LS, Mars M-69, Mars M-71, Mars M-73, Soyuz 7K-L1 , Soyuz 7K-L1E, Soyuz 7K-L1P , Venera 4V-1, DLB Beacon Lander. Liftoff Thrust: 8,847.000 kN (1,988,884 lbf). Total Mass: 707,170 kg (1,559,040 lb). Core Diameter: 4.15 m (13.61 ft). Total Length: 57.00 m (187.00 ft). Launch Price $: 70.000 million. in: 1994 price dollars.
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Model: Proton-K/D. Family: UR. Country: Russia.
Production version.
LEO Payload: 5,000 kg (11,000 lb). Payload: 4,350 kg (9,590 lb). to a: geosynchronous transfer orbit trajectory. Apogee: 400,000 km (240,000 mi). Liftoff Thrust: 9,500.000 kN (2,135,600 lbf). Total Mass: 691,500 kg (1,524,400 lb). Core Diameter: 7.40 m (24.20 ft). Total Length: 55.40 m (181.70 ft).
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Model: UR-500K/Blok D. Family: UR. Country: Russia.
Prototype version.
LEO Payload: 5,000 kg (11,000 lb). Payload: 4,350 kg (9,590 lb). to a: geosynchronous transfer orbit trajectory. Apogee: 400,000 km (240,000 mi). Liftoff Thrust: 9,500.000 kN (2,135,600 lbf). Total Mass: 691,500 kg (1,524,400 lb). Core Diameter: 7.40 m (24.20 ft). Total Length: 55.40 m (181.70 ft).
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Stage Data - Proton 8K82K / 11S824
Stage Number: 1. 1 x Stage: Proton K-1. Gross Mass: 450,510 kg (993,200 lb). Empty Mass: 31,100 kg (68,500 lb). Thrust (vac): 10,470.158 kN (2,353,785 lbf). Isp: 316 sec. Burn time: 124 sec. Isp(sl): 267 sec. Diameter: 4.15 m (13.61 ft). Span: 7.40 m (24.20 ft). Length: 21.20 m (69.50 ft). Propellants: N2O4/UDMH. No Engines: 6. Engine: RD-253-11D48. Other designations: 8S810K.
Stage Number: 2. 1 x Stage: Proton K-2. Gross Mass: 167,828 kg (369,997 lb). Empty Mass: 11,715 kg (25,827 lb). Thrust (vac): 2,399.216 kN (539,365 lbf). Isp: 327 sec. Burn time: 206 sec. Isp(sl): 230 sec. Diameter: 4.15 m (13.61 ft). Span: 4.15 m (13.61 ft). Length: 14.00 m (45.00 ft). Propellants: N2O4/UDMH. No Engines: 4. Engine: RD-0210. Other designations: 8S811K.
Stage Number: 3. 1 x Stage: Proton K-3. Gross Mass: 50,747 kg (111,877 lb). Empty Mass: 4,185 kg (9,226 lb). Thrust (vac): 630.170 kN (141,668 lbf). Isp: 325 sec. Burn time: 238 sec. Isp(sl): 230 sec. Diameter: 4.15 m (13.61 ft). Span: 4.15 m (13.61 ft). Length: 6.50 m (21.30 ft). Propellants: N2O4/UDMH. No Engines: 1. Engine: RD-0212.
Stage Number: 4. 1 x Stage: Proton 11S824. Gross Mass: 13,360 kg (29,450 lb). Empty Mass: 1,800 kg (3,900 lb). Thrust (vac): 83.300 kN (18,727 lbf). Isp: 346 sec. Burn time: 470 sec. Diameter: 3.70 m (12.10 ft). Span: 3.70 m (12.10 ft). Length: 5.50 m (18.00 ft). Propellants: Lox/Kerosene. No Engines: 1. Engine: RD-58. Other designations: 11S824; Block D; D-1-e. Block D, article number 11S824. Without guidance unit (navigation commands come from payload). Originally designed as N1-L3 lunar expedition launch vehicle lunar orbit insertion/lunar crasher stage. Adapted for use with Proton UR-500K as a fourth stage for manned circumlunar flight. Further used to launch large Lavochkin bureau unmanned lunar/planetary spacecraft. Flown from 1967 to 1975.
这是当年苏联准备用于载人环月飞行的质子火箭,但最终由于成功率太低,把握不大,又放弃了.这一放弃,苏联人就彻底与月球失之交臂了.
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在<苏联载人登月计划失败流产警示录>一文中我就狂批苏联政府没有基于质子火箭按切洛梅的设想而研制出一种起飞推力达1500吨以上的小巨型火箭,---------------------------------------------------------------------------------------------------------苏联要是按照你的想法搞小巨型火箭的话，今天他就的被别人狂批啦！！！;P
One of the many tests carried out prior to today's hot-firing test
Vinci engine hot-firing test a success
20 May 2005
Today the new Vinci cryogenic upper-stage engine was successfully ignited and fired for the first time. This marks a further milestone in developing a more efficient cryogenic upper-stage engine for future upgrades of the Ariane-5 launcher.
Although the test lasted just one second, engineers declared themselves to be “very satisfied with the test-data”. The test results will be evaluated during the next few weeks prior to a decision being made to gradually increase the duration of the test, in a step-by-step process.
Tension ran high among the engineers when the Vinci engine fired, and the hydrogen and oxygen valves opened in sequence for the first time. As the explosive mixture of hydrogen and oxygen ignited, the two turbopumps spun up to speed. Measurements confirmed after shutdown showed that the predicted values had been reached.
The Vinci engine was to have been developed under the Ariane-5 Plus Programme in combination with the development of a new Ariane-5 upper-stage, the ESC-B. Although development of the upper-stage has been put on hold after the initiation of the Ariane-5 recovery plan, the technologies involved in developing this new upper-stage engine are of importance for future developments in the European propulsion sector. As a result, it was agreed to continue a small series of test firings.
Vinci engine thrust chamber
Vinci is a re-ignitable cryogenic upper-stage engine with an expander cycle and does not require a gas generator to drive the two turbo-pumps: one for the liquid hydrogen (LH2) and one for the liquid oxygen (LOX). It will provide 18 tonnes of vacuum thrust with a specific engine impulse of 465 sec.
The successful test firing is the result of work carried out by European industry under the leadership of the prime-contractor SNECMA, and DLR’s new test facilities in Lampoldshausen, Germany. To perform hot-firing tests under near-realistic space conditions, DLR have built a new test-facility, called the P4.1 test stand. This is unique in Europe both for its size and because it allows continuous hot-firing tests of upper-stage engines, capable of providing up to 20 tonnes of thrust at pressure levels below 200 mbar (typically 60 mbar), to be carried out. The maximum test duration is 10 minutes.
The new test facility at Lampoldshausen performed flawless during today’s Vinci hot-firing test. Figure 1: The Vinci thrust chamber, developed by EADS Ottobrunn Figure 2: Vinci thrust chamber test at nominal operating point (63 bar)
毛毛斯基兄此言差矣,大型低温上面级发动机的研制可远远不是人们想象中那么容易的,这是ESA上面关于芬奇的信息.ESA可是将芬奇发动机的点火试验形容为里程碑的成就.
Vinci engine hot-firing test a success
20 May 2005
Today the new Vinci cryogenic upper-stage engine was successfully ignited and fired for the first time. This marks a further milestone in developing a more efficient cryogenic upper-stage engine for future upgrades of the Ariane-5 launcher.
Although the test lasted just one second, engineers declared themselves to be “very satisfied with the test-data”. The test results will be evaluated during the next few weeks prior to a decision being made to gradually increase the duration of the test, in a step-by-step process.
Tension ran high among the engineers when the Vinci engine fired, and the hydrogen and oxygen valves opened in sequence for the first time. As the explosive mixture of hydrogen and oxygen ignited, the two turbopumps spun up to speed. Measurements confirmed after shutdown showed that the predicted values had been reached.
The Vinci engine was to have been developed under the Ariane-5 Plus Programme in combination with the development of a new Ariane-5 upper-stage, the ESC-B. Although development of the upper-stage has been put on hold after the initiation of the Ariane-5 recovery plan, the technologies involved in developing this new upper-stage engine are of importance for future developments in the European propulsion sector. As a result, it was agreed to continue a small series of test firings.
Vinci engine thrust chamber
Vinci is a re-ignitable cryogenic upper-stage engine with an expander cycle and does not require a gas generator to drive the two turbo-pumps: one for the liquid hydrogen (LH2) and one for the liquid oxygen (LOX). It will provide 18 tonnes of vacuum thrust with a specific engine impulse of 465 sec.
The successful test firing is the result of work carried out by European industry under the leadership of the prime-contractor SNECMA, and DLR’s new test facilities in Lampoldshausen, Germany. To perform hot-firing tests under near-realistic space conditions, DLR have built a new test-facility, called the P4.1 test stand. This is unique in Europe both for its size and because it allows continuous hot-firing tests of upper-stage engines, capable of providing up to 20 tonnes of thrust at pressure levels below 200 mbar (typically 60 mbar), to be carried out. The maximum test duration is 10 minutes.
The new test facility at Lampoldshausen performed flawless during today’s Vinci hot-firing test. Figure 1: The Vinci thrust chamber, developed by EADS Ottobrunn Figure 2: Vinci thrust chamber test at nominal operating point (63 bar)
毛毛斯基兄此言差矣,大型低温上面级发动机的研制可远远不是人们想象中那么容易的,这是ESA上面关于芬奇的信息.ESA可是将芬奇发动机的点火试验形容为里程碑的成就.
Thus, with a scaled-down experiment package and with other compromises, such as the use of an elliptical orbit compared with VOIR's circular one, the Venus mission was on track again with a launch planned for May 1988.
The Challenger disaster in 1986 caused another delay. The explosion led to the reevaluation and subsequent cancellation of the Centaur G-Prime booster as cargo on the shuttle. The most powerful upper stage ever designed, Centaur was to have propelled Magellan to Venus. Its explosive liquid-oxygen and liquid-hydrogen propellants, however, were deemed too dangerous to be carried in a manned space vehicle.
The U.S. Air Force's less-powerful Inertial Upper Stage (IUS) replaced Centaur as the booster for Magellan; this required some modification of the spacecraft designs and mission plans. The aluminum Centaur-adapter structure was replaced with a lighter, graphite-epoxy frame for the IUS. A lighter spring mechanism was also used to separate the less-massive IUS from the spacecraft after burnout.
The launch procedure was changed to deploy the solar arrays before ignition of the IUS because the booster's roll-control thrusters were too close to the ends of the solar panels while in their stowed (folded) position. Lastly, rather than subject the entire spacecraft to a repetition of full static tests in the new IUS configuration, a mockup Magellan structure was used. Fidelity was assured by using real components borrowed from a Voyager spacecraft on public display at the National Air and Space Museum in Washington, D.C.
The loss of the Challenger and the 32-month suspension of shuttle missions delayed and reshuffled many planned space activities. The Galileo mission to Jupiter, for one, would have to launch in October 1989, the date initially set for Magellan, or wait another two years for the necessary alignment of planets. The result for Magellan was an early May 1989 launch and the use of a Type-IV trajectory. This meant that the spacecraft would spend 15 months traveling one and a half times around the Sun before arriving at Venus. The original May 1988 launch date would have allowed Magellan to reach Venus in 4 months by traveling less than 180 degrees around the Sun via a Type-I trajectory.
Thus, the $551 million mission (see Table 2-1) and the spacecraft that will soon arrive at Venus are much different than NASA had planned a decade earlier, yet the basic scientific mapping objectives remain unchanged
这是从JPL上得到的资料,当年挑战者航天飞机爆炸之后美国为了更安全,在发射麦哲伦金星探测器时就将原计划利用的半人马座低温液体上面级火箭改为使用固体燃料上面级火箭.麦哲伦可是大型星际飞船,由此可见美国固体燃料上面级火箭技术的强悍.事实上只使到今天为止,国内网友们包括我本人在内都对发射星际飞船所必需的上面级(地球脱离级)技术知道堪少,期望高手们详细科普.
The Challenger disaster in 1986 caused another delay. The explosion led to the reevaluation and subsequent cancellation of the Centaur G-Prime booster as cargo on the shuttle. The most powerful upper stage ever designed, Centaur was to have propelled Magellan to Venus. Its explosive liquid-oxygen and liquid-hydrogen propellants, however, were deemed too dangerous to be carried in a manned space vehicle.
The U.S. Air Force's less-powerful Inertial Upper Stage (IUS) replaced Centaur as the booster for Magellan; this required some modification of the spacecraft designs and mission plans. The aluminum Centaur-adapter structure was replaced with a lighter, graphite-epoxy frame for the IUS. A lighter spring mechanism was also used to separate the less-massive IUS from the spacecraft after burnout.
The launch procedure was changed to deploy the solar arrays before ignition of the IUS because the booster's roll-control thrusters were too close to the ends of the solar panels while in their stowed (folded) position. Lastly, rather than subject the entire spacecraft to a repetition of full static tests in the new IUS configuration, a mockup Magellan structure was used. Fidelity was assured by using real components borrowed from a Voyager spacecraft on public display at the National Air and Space Museum in Washington, D.C.
The loss of the Challenger and the 32-month suspension of shuttle missions delayed and reshuffled many planned space activities. The Galileo mission to Jupiter, for one, would have to launch in October 1989, the date initially set for Magellan, or wait another two years for the necessary alignment of planets. The result for Magellan was an early May 1989 launch and the use of a Type-IV trajectory. This meant that the spacecraft would spend 15 months traveling one and a half times around the Sun before arriving at Venus. The original May 1988 launch date would have allowed Magellan to reach Venus in 4 months by traveling less than 180 degrees around the Sun via a Type-I trajectory.
Thus, the $551 million mission (see Table 2-1) and the spacecraft that will soon arrive at Venus are much different than NASA had planned a decade earlier, yet the basic scientific mapping objectives remain unchanged
这是从JPL上得到的资料,当年挑战者航天飞机爆炸之后美国为了更安全,在发射麦哲伦金星探测器时就将原计划利用的半人马座低温液体上面级火箭改为使用固体燃料上面级火箭.麦哲伦可是大型星际飞船,由此可见美国固体燃料上面级火箭技术的强悍.事实上只使到今天为止,国内网友们包括我本人在内都对发射星际飞船所必需的上面级(地球脱离级)技术知道堪少,期望高手们详细科普.
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支持你 "高凉陈君CT"，在大本营看了这么长时间，这里充满了不屑一顾的回帖，很少有你这样
认证查找资料进行分析的楼主，支持你，什么事情要有个谦虚的态度，鄙视 毛毛斯基 之流的人
如果是高手，不会对别人的努力不屑一顾的。动不动回几个字，好像自己水平多高一样。
认证查找资料进行分析的楼主，支持你，什么事情要有个谦虚的态度，鄙视 毛毛斯基 之流的人
如果是高手，不会对别人的努力不屑一顾的。动不动回几个字，好像自己水平多高一样。
原帖由 高凉陈君CT 于 2007-12-31 04:01 发表
The Inertial Upper Stage or IUS is a two-stage solid-fueled booster rocket developed by NASA and the U.S. Air Force for the launching of large payloads from either a Titan III (later Titan IV) rocket ...
事实上毛人要登月的话根本不用什么低温上面级，能源两次发射就够了，能源的货物自带的发动机也用不着氢氧的。40-50吨的氢氧对毛国来说毫无用处。
原帖由 mcuttle 于 2007-12-31 09:26 发表
支持你 "高凉陈君CT"，在大本营看了这么长时间，这里充满了不屑一顾的回帖，很少有你这样
认证查找资料进行分析的楼主，支持你，什么事情要有个谦虚的态度，鄙视 毛毛斯基 之流的人
如果是高手，不会对别人的努力不 ...
不支持楼主的又不是偶一人，CD坛上从版主贵宾到偶没几个支持的，难道这些人都错了！！;P
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Current Missions - Cassini
Spacecraft
Launch: October 15, 1997
Mass: 5,712 kilograms (12,593 pounds), consisting of 2,125-kilogram (4,685-pound) orbiter, 320-kilogram (705-pound) Huygens probe, launch vehicle adapter and 3,132 kilograms (6,905 pounds) of propellants
Science instruments: Orbiter optical camera system, imaging radar, radio science, ion and neutral mass spectrometer, visible and infrared mapping spectrometer, composite infrared spectrometer, cosmic dust analyzer, radio and plasma wave spectrometer, plasma spectrometer, ultraviolet imaging spectrograph, magnetospheric imaging instrument, dual technique magnetometer; Huygens probe descent imager and spectral radiometer, atmospheric structure instrument, gas chromatograph and mass spectrometer, aerosol collector pyrolyzer, surface science package, doppler wind experiment
Overview
The Cassini mission to Saturn is the most ambitious effort in planetary space exploration ever mounted. A joint endeavor of NASA, the European Space Agency (ESA) and the Italian Space Agency (known as ASI for its acronym in Italian), Cassini is sending a sophisticated robotic spacecraft to orbit the ringed planet and study the Saturnian system in detail over a four-year period.
Onboard Cassini is a scientific probe called Huygens that will be released from the main spacecraft to parachute through the atmosphere to the surface of Saturn's largest and most interesting moon, Titan, which is shrouded by an opaque atmosphere. Titan's atmosphere includes organic compounds leading scientists to believe that the moon may be like a frozen vault of conditions similar to those on Earth before life began. The Cassini orbiter will also use imaging radar to map Titan's surface.
Launched October 15, 1997, on a Titan 4 rocket from Cape Canaveral, Florida, Cassini has flown past other planets on its way to Saturn -- once each by Earth and Jupiter, twice by Venus -- to borrow gravitational energy to speed it on its way. Cassini will enter Saturn orbit July 1, 2004, and the Huygens probe will descend to the surface of Titan on January 14, 2005. More Information:
这是美国的卡西尼土星探测器,重达5.7吨.用大力神四系列火箭发射.而神舟飞船重也不过7.8吨,用比大力神四火箭大一点的火箭就足以将其送入月球轨道.而美国的大力神四B系列火箭起飞质量都在929至940吨之间.而中国未来1440吨级起飞推力的长征五号火箭的起飞质量都会远远超过1000吨.
事实上中国要实现载人环月飞行真的已经不存在任何不可克服的技术阻碍.而中国政府目前对这一计划的兴趣也越来越浓厚.估计在未来四年内就应该有环月飞行的具体计划公布,毕竞CEO的任期有限啊.而且又不必投入太多的(无论是火箭还是飞船都已经相当成熟与有把握)金钱就可以达到历史留名的效果,这种成就对政治领导人而言是很具透惑力的.这可是中国漫长历史过程中早已经演化成熟的一条不成文潜规则.
Spacecraft
Launch: October 15, 1997
Mass: 5,712 kilograms (12,593 pounds), consisting of 2,125-kilogram (4,685-pound) orbiter, 320-kilogram (705-pound) Huygens probe, launch vehicle adapter and 3,132 kilograms (6,905 pounds) of propellants
Science instruments: Orbiter optical camera system, imaging radar, radio science, ion and neutral mass spectrometer, visible and infrared mapping spectrometer, composite infrared spectrometer, cosmic dust analyzer, radio and plasma wave spectrometer, plasma spectrometer, ultraviolet imaging spectrograph, magnetospheric imaging instrument, dual technique magnetometer; Huygens probe descent imager and spectral radiometer, atmospheric structure instrument, gas chromatograph and mass spectrometer, aerosol collector pyrolyzer, surface science package, doppler wind experiment
Overview
The Cassini mission to Saturn is the most ambitious effort in planetary space exploration ever mounted. A joint endeavor of NASA, the European Space Agency (ESA) and the Italian Space Agency (known as ASI for its acronym in Italian), Cassini is sending a sophisticated robotic spacecraft to orbit the ringed planet and study the Saturnian system in detail over a four-year period.
Onboard Cassini is a scientific probe called Huygens that will be released from the main spacecraft to parachute through the atmosphere to the surface of Saturn's largest and most interesting moon, Titan, which is shrouded by an opaque atmosphere. Titan's atmosphere includes organic compounds leading scientists to believe that the moon may be like a frozen vault of conditions similar to those on Earth before life began. The Cassini orbiter will also use imaging radar to map Titan's surface.
Launched October 15, 1997, on a Titan 4 rocket from Cape Canaveral, Florida, Cassini has flown past other planets on its way to Saturn -- once each by Earth and Jupiter, twice by Venus -- to borrow gravitational energy to speed it on its way. Cassini will enter Saturn orbit July 1, 2004, and the Huygens probe will descend to the surface of Titan on January 14, 2005. More Information:
这是美国的卡西尼土星探测器,重达5.7吨.用大力神四系列火箭发射.而神舟飞船重也不过7.8吨,用比大力神四火箭大一点的火箭就足以将其送入月球轨道.而美国的大力神四B系列火箭起飞质量都在929至940吨之间.而中国未来1440吨级起飞推力的长征五号火箭的起飞质量都会远远超过1000吨.
事实上中国要实现载人环月飞行真的已经不存在任何不可克服的技术阻碍.而中国政府目前对这一计划的兴趣也越来越浓厚.估计在未来四年内就应该有环月飞行的具体计划公布,毕竞CEO的任期有限啊.而且又不必投入太多的(无论是火箭还是飞船都已经相当成熟与有把握)金钱就可以达到历史留名的效果,这种成就对政治领导人而言是很具透惑力的.这可是中国漫长历史过程中早已经演化成熟的一条不成文潜规则.
问几个问题：
地球转移轨道的CE1，仅重2吨多。应该是长征最大的转移轨道载荷了吧？
这个长三C没有用助推火箭，如果采用，对转移轨道的载荷有多大影响？
另外，载人直接转移的轨道方式和CE1的微调转移方式应该在其火箭载荷能力方面有很大区别吧？
地球转移轨道的CE1，仅重2吨多。应该是长征最大的转移轨道载荷了吧？
这个长三C没有用助推火箭，如果采用，对转移轨道的载荷有多大影响？
另外，载人直接转移的轨道方式和CE1的微调转移方式应该在其火箭载荷能力方面有很大区别吧？
看不懂啊，
不过对高手总是要顶一顶的。。。
不过对高手总是要顶一顶的。。。
氢氧煤油固推都要搞,而中国人的传统又是要一个大项目来带若干个小项目否则就搞不好,长5就是这个大项目,长5成熟以后不管是360吨煤油还是200吨氢氧都有希望.至于是用来登月还是干别的什么就不那么重要了.
原帖由 tomluter 于 2007-12-31 15:11 发表
问几个问题：
地球转移轨道的CE1，仅重2吨多。应该是长征最大的转移轨道载荷了吧？
这个长三C没有用助推火箭，如果采用，对转移轨道的载荷有多大影响？
另外，载人直接转移的轨道方式和CE1的微调转移方式应该在其 ...
嫦娥是用长3甲打的,长三丙有2个助推器.
带助推以及数量对GEO轨道运载能力的影响可参考下面几个图.
长征三乙现在最大的GTO是5吨1，长三乙地月转移轨道是4吨以上，够嫦娥第二阶段完成落的任务。
TG的目光就是远大啊:victory:
原帖由 重剑无锋 于 2008-1-1 13:45 发表
长征三乙现在最大的GTO是5吨1，长三乙地月转移轨道是4吨以上，够嫦娥第二阶段完成落的任务。
长三乙地月转移轨道运载能力是3.4吨,未来长5最大构型10.6吨.
水平不够,就不出来现丑了,这样的帖子要 顶
貌似有些道理，顶一个
原帖由 ddeell72 于 2008-1-1 16:50 发表
我个人认为空射航天器值得努努力搞一下：先不弄太复杂的，就空射火箭。
目前的火箭，一半左右的燃料、推力仅仅为了离开地面，我觉得太浪费。国家在搞大飞机、06年珠海航展展出过轰六空射小型火箭发射小卫星的方案 ...
空射的优点是很明显的。比如美国的飞马座通过带小翼优化飞行弹道等措施，空射的运载能力比类似构形在地面发射大了1倍。但是受的制约也是很明显的，就是受载机限制，很难做太大。所以这个我们是要做的，但是只是整个计划的一个组成部分。
至于重型火箭，除了登月，几乎看不到其他用处。
:lol
老兄好久不见了,新年快乐哦!:handshake :lol
原帖由 shh 于 2008-1-2 15:08 发表
空射的优点是很明显的。比如美国的飞马座通过带小翼优化飞行弹道等措施，空射的运载能力比类似构形在地面发射大了1倍。但是受的制约也是很明显的，就是受载机限制，很难做太大。所以这个我们是要做的，但是只是整 ...
老兄好久不见了,新年快乐哦!:handshake :lol
原帖由 毛毛斯基 于 2007-12-31 10:07 发表
不支持楼主的又不是偶一人，CD坛上从版主贵宾到偶没几个支持的，难道这些人都错了！！;P
不是这些人都错了，只是大家应该有个讨论的分为，我来这个论坛一个多月了，看到了80%的回帖都是三言两语感觉自己多高深，
没有几个像楼主这样认真分析问题的，错误在所难免，关键在分析，而不是懒得说话，懒得分析。我也是做工程的，反正我周围
越是高手越谦虚，越分析的周到细致。其余的只能说是伪高手，或者只能说水平还有，但是到不了一定境界，到了一定境界的
给人感觉完全不同，在这里真的很少。
原帖由 hbao 于 2008-1-1 18:15 发表
长三乙地月转移轨道运载能力是3.4吨,未来长5最大构型10.6吨.
才3点4？这个是月球轨道能力了吧，不是转移轨道？
原帖由 重剑无锋 于 2008-1-2 20:21 发表
才3点4？这个是月球轨道能力了吧，不是转移轨道？
3.4吨不小了
就是地月转移轨道.
想对应的月球软着陆运载能力是1.25吨,应该足够了.
原帖由 shh 于 2008-1-2 15:08 发表
空射的优点是很明显的。比如美国的飞马座通过带小翼优化飞行弹道等措施，空射的运载能力比类似构形在地面发射大了1倍。但是受的制约也是很明显的，就是受载机限制，很难做太大。所以这个我们是要做的，但是只是整 ...
我一直觉得那个“神龙”就是我们的飞马座……
空射如果用飞艇会怎么样?毕竟飞机那点初速也没多大意义.
楼主精神值得赞赏，但似乎考虑问题也有点过于集中于某一点，尤其是分析问题总喜欢计算起飞推力大小，而没有更多考虑火箭发射全过程的情况。俺估计还是那个登月情节引起的。因为登月需要大推力火箭，一个衡量标准就是起飞推力有多大。但只关注起飞推力，不去了解整个火箭发射过程，分析问题会有所偏颇。
从航天飞行效益来讲，高度越高，速度越快，付出的代价越大，这就使得在上面级上，采用氢氧发动机更合适，打个极端的比方，如果上面级用的是个黑火药火箭，估计会被人骂死，这足可以看作是用奔驰轿车在运煤炭差不多。所以上面级要用高比冲的东西。而目前所知的高比冲发动机，就是氢氧发动机。所以，没有采用都是煤油发动机方案，这方面是一个很重要的原因。
全煤油或者三级的上面级采用现有氢氧发动机的整体方案，所用的发动机数量，运载能力上，比现有的方案的总发射成本上，任务灵活性上也不具有优势。龙乐豪的文章上，关于这两者的比较，写的还是比较明白的。原因在于芯级采用氢氧发动机，可以实现所谓一级半到低轨道，二级半到高轨道方案，同时也减少发动机数量，从而降低整个发射成本，提高运载效率。
可以说，现在航天发射，几乎都采用了所谓捆绑的一级半或者二级半策略，包括俄罗斯的能源号和美国最新的登月火箭战神五、欧洲的阿五、日本的H2A、我国的长五。这是航天发射技术发展到今天大家选择的相对合适的火箭构造技术方案。
对于氢氧发动机研制，龙乐豪的文章写得很清楚，即未来上面级要用的到，当然以后用于登月也是可能的。不过，真的要登月的话，在运载火箭层面，还是大推力火箭发动机的研制。置于是360吨还是500吨，或者如一些网友猜测的是四120吨推力共用燃烧室的480吨-500吨并联发动机之类，航天专家都会算排列组合的帐，来看最终选择什么样的推力进行研制最合适。
最终，俺感觉楼主是化了心思的。不过，建议暂时将登月放在一边，将目光放在航天应用，或者就是火箭本身，放在目前主要的集中火箭的构型，火箭运载能力组合方面的优缺点比较上，或许会有更多更全面的收获，然后在此基础上再去思考登月该选择什么样的技术，会有更进一步的认识。如果感兴趣，可以到维普网，这个网是一个关于期刊和学术论文的网站，需要花钱下载一些文章，包括龙乐豪关于未来航天发展的文章，其中也包括长五和登月火箭研制的思考等。
从航天飞行效益来讲，高度越高，速度越快，付出的代价越大，这就使得在上面级上，采用氢氧发动机更合适，打个极端的比方，如果上面级用的是个黑火药火箭，估计会被人骂死，这足可以看作是用奔驰轿车在运煤炭差不多。所以上面级要用高比冲的东西。而目前所知的高比冲发动机，就是氢氧发动机。所以，没有采用都是煤油发动机方案，这方面是一个很重要的原因。
全煤油或者三级的上面级采用现有氢氧发动机的整体方案，所用的发动机数量，运载能力上，比现有的方案的总发射成本上，任务灵活性上也不具有优势。龙乐豪的文章上，关于这两者的比较，写的还是比较明白的。原因在于芯级采用氢氧发动机，可以实现所谓一级半到低轨道，二级半到高轨道方案，同时也减少发动机数量，从而降低整个发射成本，提高运载效率。
可以说，现在航天发射，几乎都采用了所谓捆绑的一级半或者二级半策略，包括俄罗斯的能源号和美国最新的登月火箭战神五、欧洲的阿五、日本的H2A、我国的长五。这是航天发射技术发展到今天大家选择的相对合适的火箭构造技术方案。
对于氢氧发动机研制，龙乐豪的文章写得很清楚，即未来上面级要用的到，当然以后用于登月也是可能的。不过，真的要登月的话，在运载火箭层面，还是大推力火箭发动机的研制。置于是360吨还是500吨，或者如一些网友猜测的是四120吨推力共用燃烧室的480吨-500吨并联发动机之类，航天专家都会算排列组合的帐，来看最终选择什么样的推力进行研制最合适。
最终，俺感觉楼主是化了心思的。不过，建议暂时将登月放在一边，将目光放在航天应用，或者就是火箭本身，放在目前主要的集中火箭的构型，火箭运载能力组合方面的优缺点比较上，或许会有更多更全面的收获，然后在此基础上再去思考登月该选择什么样的技术，会有更进一步的认识。如果感兴趣，可以到维普网，这个网是一个关于期刊和学术论文的网站，需要花钱下载一些文章，包括龙乐豪关于未来航天发展的文章，其中也包括长五和登月火箭研制的思考等。
有资料上说长三乙地月转移轨道运载能力是3.2吨，而嫦娥二期的载荷是3.65吨。所有要对火箭进行改进。
龙老的文章里在比较只采用一种发动机方案时算只采用液氧煤油发动机的一级半方案起飞质量1600吨，要比现有方案大一倍以上。实际上这个是一个极端。完全可以采用2级半的方案。类似的例子就是Atlas V的重型版本，芯一级加助推级3台RD-180，上面级用1台或2台改进的半人马座。起飞质量近1000吨，运载能力LEO25吨/GTO13吨。这个对应于我们的方案差不多就是芯级和助推器用大约10台YF-100 120吨级发动机，上面级用一个改进的YF-75。从运载能力讲，这个或许也是可以的。
但是CZ-5作为我国运载工具以后几十年的主力(也就是说难得一上的重大型号)，其研制必须从大局考虑，其带动的技术发展要符合以后一段时间里我国航天发展的需求。从这个角度讲，研制一型大推力氢氧发动机是必要的。登月会是一个考虑，但是并不是唯一一个考虑。在长征5研制完成之后，我国大中型航天发射载具的发展一个是重型火箭(类似于登月这样的任务)，一个是可复用发射载具。(其他还有一些诸如空射火箭，小型火箭，先进上面级等，但是涉及的发动机推力都不会很大)。在这两个方向上，先进的大推力氢氧发动机都是非常可能被采用的技术。考虑到在这以前我国仅研制过小推力的YF-73/75氢氧发动机，因此CZ-5项目有必要带动这个方向的发展，通过完成一个大推力氢氧发动机型号研制的全过程，为以后这个方向上的发展打下基础。
凡事预则立。象登月这么大的项目，没有一些提前考虑是不行的。但是有一些提前考虑，也并不意味着我国将来就一定要实施这个项目。登月从来就是政治项目，这是一个政治决定，会由当时的很多情况来综合决定。就象航天界已经在公开讨论重型火箭的研制问题，作为航天部门，必要的准备是应该的，但是并不是说我国载人登月项目已经开始实施。
另外，CZ-5里的5m构形与3.35m构形在起飞推力类似情况下，运载能力还是有些区别的。毕竟氢氧发动机的比冲高不少。
但是CZ-5作为我国运载工具以后几十年的主力(也就是说难得一上的重大型号)，其研制必须从大局考虑，其带动的技术发展要符合以后一段时间里我国航天发展的需求。从这个角度讲，研制一型大推力氢氧发动机是必要的。登月会是一个考虑，但是并不是唯一一个考虑。在长征5研制完成之后，我国大中型航天发射载具的发展一个是重型火箭(类似于登月这样的任务)，一个是可复用发射载具。(其他还有一些诸如空射火箭，小型火箭，先进上面级等，但是涉及的发动机推力都不会很大)。在这两个方向上，先进的大推力氢氧发动机都是非常可能被采用的技术。考虑到在这以前我国仅研制过小推力的YF-73/75氢氧发动机，因此CZ-5项目有必要带动这个方向的发展，通过完成一个大推力氢氧发动机型号研制的全过程，为以后这个方向上的发展打下基础。
凡事预则立。象登月这么大的项目，没有一些提前考虑是不行的。但是有一些提前考虑，也并不意味着我国将来就一定要实施这个项目。登月从来就是政治项目，这是一个政治决定，会由当时的很多情况来综合决定。就象航天界已经在公开讨论重型火箭的研制问题，作为航天部门，必要的准备是应该的，但是并不是说我国载人登月项目已经开始实施。
另外，CZ-5里的5m构形与3.35m构形在起飞推力类似情况下，运载能力还是有些区别的。毕竟氢氧发动机的比冲高不少。
从火箭的成本来讲，发动机要占到相当一部分。通过二级半结构达到一级半结构达到的效果，从理论上、实践上都可行，但经济上是否有优势，就不好说了。
当然，航天研发要有前瞻性，这是应该的。正好龙他们是这样的人。
当然，航天研发要有前瞻性，这是应该的。正好龙他们是这样的人。
原帖由 TSQ 于 2008-1-3 15:17 发表
如果感兴趣，可以到维普网，这个网是一个关于期刊和学术论文的网站，需要花钱下载一些文章，包括龙乐豪关于未来航天发展的文章，其中也包括长五和登月火箭研制的思考等。
就算是没有读过大学的人，用google和baidu搜搜也经常会有维普等几个期刊数据库
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支持楼主的研究态度，真的好认真，好深入。
不过楼主关于‘苏联没有发射绕月载人飞船是错误的’ 这一论点本人不大认同， 当时的情况就是谁先把脚印印在月球上谁就赢， 没有必要也没有资源在中间搞一个绕月载人计划。
不过楼主关于‘苏联没有发射绕月载人飞船是错误的’ 这一论点本人不大认同， 当时的情况就是谁先把脚印印在月球上谁就赢， 没有必要也没有资源在中间搞一个绕月载人计划。