The LE-7 and its succeeding upgrade model the LE-7A are staged combustion cycleLH2/LOXliquid rocket engines produced in Japan for the H-II series of launch vehicles. Design and production work was all done domestically in Japan, the first major (main/first-stage) liquid rocket engine with that claim, in a collaborative effort from the National Space Development Agency (NASDA), Aerospace Engineering Laboratory (NAL), Mitsubishi Heavy Industries, and Ishikawajima-Harima. NASDA and NAL have since been integrated into JAXA. However, a large part of the work was contracted to Mitsubishi, with Ishikawajima-Harima providing turbomachinery, and the engine is often referred to as the Mitsubishi LE-7(A).
The fuel turbopump had an issue using the originally designed inducer (a propeller-like axial pump used to raise the inlet pressure of the propellant ahead of the main turbopumps to prevent cavitation) where the inducer would itself begin to cavitate and cause an imbalance resulting in excessive vibration. A comprehensive post-flight analysis of the unsuccessful 8th H-II launch, including a deep ocean retrieval of the wreckage, determined that fatigue due to this vibration was the cause of premature engine failure.
LE-7Aedit
The LE-7A is an upgraded model from the LE-7 rocket engine. Basic design is unchanged from the original model. The 7A had additional engineering effort placed on cost cutting, reliability, and performance developments. The renovation was undertaken to mate it with the likewise improved H-IIA launch vehicle, with the common goal being a more reliable, more powerful and flexible, and more cost effective launch system.
Changes / improvementsedit
Specific emphasis was placed on reducing or the amount of required welding by allowing for more machined or cast components, and to simplify as many of the remaining welds as possible. This resulted in a substantial rework of the pipe routing (which makes the outward appearance of the two models considerably different). To combat the fuel inducer complications described above, the fuel inducer was redesigned for the 7A. The oxidizer inducer was also redesigned, but this was primarily due to poor performance at low inlet pressures as opposed to reliability concerns. The fuel turbopump itself was also the subject of various durability enhancements. Additionally the combustion chamber/injector assembly underwent a number of small changes, like decreasing the number of injector elements, to reduce machining complexity (and thus cost) and improve reliability. While these changes overall resulted in a drop in maximum specific impulse to 440 seconds (4.3 km/s) (basically making the engine less fuel efficient), the trade off for lower cost and enhanced reliability was considered acceptable.
New nozzle design (side-loading problem)edit
For the new engine model, a nozzle extension was designed that could be added to the base of the new standard “short” nozzle when extra performance was required. But when the engine was fitted with the nozzle extension, the 7A encountered a new problem with unprecedented side-loads and irregular heating on the nozzle strong enough to damage the gimbalactuators and regenerative cooling tubes during startup. Meticulous computational fluid dynamics (CFD) work was able to sufficiently replicate and trace the dangerous transient loading and a new one-piece “long” nozzle with full regenerative cooling (as opposed to the original short nozzle with a separate film-cooled extension) was designed to mitigate the problem. Before this new nozzle was ready, some H-IIA's were launched using only the short nozzle. The 7A no longer uses a separate nozzle extension in any configuration.[1]
Use on H-IIBedit
The new H-IIB launch vehicles uses two LE-7A engines in its first stage.[2]
LE-7A specificationsedit
Operational cycle: staged combustion
Fuel: hydrogen
Oxidizer: liquid oxygen
Mixture ratio (oxidizer to fuel): 5.90
Short nozzle:
Rated thrust (sea level): 843 kN (190,000 lbf)
Rated thrust (vacuum): 1,074 kN (241,000 lbf)
Specific impulse (sea level):
Specific impulse (vacuum): 429 seconds (4.21 km/s)
Long nozzle:
Rated thrust (sea level): 870 kN (200,000 lbf)
Rated thrust (vacuum): 1,098 kN (247,000 lbf)
Specific impulse (sea level): 338 seconds (3.31 km/s)
^Yasuhide Watanabe; Norio Sakazume; Masanori Tsuboi (July 2003). "LE-7A Engine Separation Phenomenon Differences of the Two Nozzle Configurations". 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit. doi:10.2514/6.2003-4763. ISBN978-1-62410-098-7. Retrieved 13 May 2023.
^"Overview of the H-IIB Launch Vehicle". JAXA. Retrieved 13 May 2023.
External linksedit
Wikimedia Commons has media related to LE-7.
Japanese Wikipedia LE-7 page (in Japanese)
Japanese Wikipedia LE-7A page (in Japanese)
H-IIA Rocket Engine Development
Overview of the H-IIA Launch Vehicle Includes H-IIA-F6 failure analysis
December 15, 2023
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The LE 7 and its succeeding upgrade model the LE 7A are staged combustion cycle LH2 LOX liquid rocket engines produced in Japan for the H II series of launch vehicles Design and production work was all done domestically in Japan the first major main first stage liquid rocket engine with that claim in a collaborative effort from the National Space Development Agency NASDA Aerospace Engineering Laboratory NAL Mitsubishi Heavy Industries and Ishikawajima Harima NASDA and NAL have since been integrated into JAXA However a large part of the work was contracted to Mitsubishi with Ishikawajima Harima providing turbomachinery and the engine is often referred to as the Mitsubishi LE 7 A LE 7LE 7 Nagoya City Science Museum 2006Country of originJapanDesignerJAXAManufacturerMitsubishi Heavy IndustriesApplicationBoosterStatusSucceeded by LE 7A upgradeLiquid fuel enginePropellantLOX LH2Mixture ratio5 9 1CycleStaged combustionConfigurationChamber1Nozzle ratio52 1PerformanceThrust vacuum1 078 kN 242 000 lbf Thrust sea level843 5 kN 189 600 lbf Thrust to weight ratio64 13Chamber pressure12 7 MPa 1 840 psi Specific impulse vacuum446 seconds 4 37 km s Specific impulse sea level349 seconds 3 42 km s DimensionsLength3 4 mDry weight1 714 kg 3 779 lb Used inH II first stageLE 7A Mitsubishi Heavy Industries show room Shinagawa Japan The original LE 7 was an expendable high efficiency medium sized motor with sufficient thrust for use on the H II Contents 1 H II Flight 8 only operational LE 7 failure 2 LE 7A 2 1 Changes improvements 2 2 New nozzle design side loading problem 2 3 Use on H IIB 2 4 LE 7A specifications 3 See also 4 References 5 External linksH II Flight 8 only operational LE 7 failure editThe fuel turbopump had an issue using the originally designed inducer a propeller like axial pump used to raise the inlet pressure of the propellant ahead of the main turbopumps to prevent cavitation where the inducer would itself begin to cavitate and cause an imbalance resulting in excessive vibration A comprehensive post flight analysis of the unsuccessful 8th H II launch including a deep ocean retrieval of the wreckage determined that fatigue due to this vibration was the cause of premature engine failure LE 7A editThe LE 7A is an upgraded model from the LE 7 rocket engine Basic design is unchanged from the original model The 7A had additional engineering effort placed on cost cutting reliability and performance developments The renovation was undertaken to mate it with the likewise improved H IIA launch vehicle with the common goal being a more reliable more powerful and flexible and more cost effective launch system Changes improvements edit Specific emphasis was placed on reducing or the amount of required welding by allowing for more machined or cast components and to simplify as many of the remaining welds as possible This resulted in a substantial rework of the pipe routing which makes the outward appearance of the two models considerably different To combat the fuel inducer complications described above the fuel inducer was redesigned for the 7A The oxidizer inducer was also redesigned but this was primarily due to poor performance at low inlet pressures as opposed to reliability concerns The fuel turbopump itself was also the subject of various durability enhancements Additionally the combustion chamber injector assembly underwent a number of small changes like decreasing the number of injector elements to reduce machining complexity and thus cost and improve reliability While these changes overall resulted in a drop in maximum specific impulse to 440 seconds 4 3 km s basically making the engine less fuel efficient the trade off for lower cost and enhanced reliability was considered acceptable New nozzle design side loading problem edit For the new engine model a nozzle extension was designed that could be added to the base of the new standard short nozzle when extra performance was required But when the engine was fitted with the nozzle extension the 7A encountered a new problem with unprecedented side loads and irregular heating on the nozzle strong enough to damage the gimbal actuators and regenerative cooling tubes during startup Meticulous computational fluid dynamics CFD work was able to sufficiently replicate and trace the dangerous transient loading and a new one piece long nozzle with full regenerative cooling as opposed to the original short nozzle with a separate film cooled extension was designed to mitigate the problem Before this new nozzle was ready some H IIA s were launched using only the short nozzle The 7A no longer uses a separate nozzle extension in any configuration 1 Use on H IIB edit The new H IIB launch vehicles uses two LE 7A engines in its first stage 2 LE 7A specifications edit Operational cycle staged combustion Fuel hydrogen Oxidizer liquid oxygen Mixture ratio oxidizer to fuel 5 90 Short nozzle Rated thrust sea level 843 kN 190 000 lbf Rated thrust vacuum 1 074 kN 241 000 lbf Specific impulse sea level Specific impulse vacuum 429 seconds 4 21 km s Long nozzle Rated thrust sea level 870 kN 200 000 lbf Rated thrust vacuum 1 098 kN 247 000 lbf Specific impulse sea level 338 seconds 3 31 km s Specific impulse vacuum 440 seconds 4 3 km s Dry mass 1 800 kg 4 000 lb Length short nozzle 3 2 m long nozzle 3 7 m Throttle capability 72 100 Thrust to weight 65 9 Nozzle area ratio 51 9 1 Combustion chamber pressure 12 0 MPa 1 740 psi Liquid hydrogen turbopump 41 900 rpm Liquid oxygen turbopump 18 300 rpmSee also editLE 5 LE 9 H II H IIA amp H IIB Comparison of orbital rocket engines Liquid fuel rocket Staged combustion cycle JAXAReferences edit Yasuhide Watanabe Norio Sakazume Masanori Tsuboi July 2003 LE 7A Engine Separation Phenomenon Differences of the Two Nozzle Configurations 39th AIAA ASME SAE ASEE Joint Propulsion Conference and Exhibit doi 10 2514 6 2003 4763 ISBN 978 1 62410 098 7 Retrieved 13 May 2023 Overview of the H IIB Launch Vehicle JAXA Retrieved 13 May 2023 External links edit nbsp Wikimedia Commons has media related to LE 7 Encyclopedia Astronautica info page on the LE 7 Encyclopedia Astronautica info page on the LE 7A Japanese Wikipedia LE 7 page in Japanese Japanese Wikipedia LE 7A page in Japanese H IIA Rocket Engine Development Overview of the H IIA Launch Vehicle Includes H IIA F6 failure analysis Retrieved from https en wikipedia org w index php title LE 7 amp oldid 1170775071, wikipedia, wiki, book, books, library,
