بررسی اثر ابعاد هندسی پاشنده بر عملکرد محفظه احتراق رانشگر دومولفه‌ای

نوع مقاله : مقاله پژوهشی

نویسندگان

1 دانشکده مهندسی هوافضا دانشگاه صنعتی امیرکبیر

2 دانشگاه صنعتی شریف

3 پژوهشکده سامانه های حمل و نقل فضایی

چکیده

استفاده از احتراق پیشرانه‌های خودمشتعل در رانشگرها، به ­دلیل دمای بالای محصولات احتراق، سبب افزایش ضربه ویژه می‌شود. در این مقاله، با استفاده از یک نرم‌افزار توسعه داده‌ شده، فرایند احتراق درون رانشگر دومولفه‌ای به‌صورت یک‌بعدی و با استفاده از سینتیک شیمیایی شبیه‌سازی می‌شود. در این راستا، مدل‌هایی برای پاشش، تبخیر قطرات، تشکیل فیلم مایع و محاسبات مربوط به انتقال حرارت از فیلم‌های مایع و گازی و احتراق به­ کار گرفته شده است. با استفاده از این نرم‌افزار، رفتار رانشگر آستریوم با سوخت منومتیل‌هیدرازین و اکسنده تتراکسید نیتروژن شبیه‌سازی شده است. با بهره‌گیری از مکانیزم شیمیایی گسترده 1619مرحله‌ای، نتایج شبیه‌سازی عملکرد رانشگر در دبی‌های مختلف اعتبارسنجی شده است. سپس، اثر ابعاد هندسی پاشنده بر فرایند تبخیر قطرات و نیز احتراق مورد بررسی دقیق قرار گرفته است. نتایج نشان می‌دهد که بزرگ­شدن پاشنده سبب افزایش طول تبخیر قطرات شده و ساختار شعله درون محفظه احتراق تغییر می‌کند، به ­نحوی که محصولات احتراق با دمای بالاتر وارد نازل شده و درنتیجه ضربه ویژه رانشگر افزایش می‌یابد. 

کلیدواژه‌ها

موضوعات

عنوان مقاله [English]

Investigation of injector dimension on the performance of combustion chamber of a bi-propelant thruster

نویسندگان [English]

  • Masoud EidiAttarZade 1
  • محمد فرشچی 2
  • Atiyeh Sarabadani 2
  • hamed khosrobeygi 2
  • ghazal davarnia 2
  • Alireza Ramezani 3
1 Aerospace Engineering Department Amirkabir University of Technology
2 Sharif University of Technology
3 Space Transportation Research Institute, Iranian Space Research Center, Tehran, Iran
چکیده [English]

Combustion of hypergolic propellants increases the specific impulse in the thrusters due to high temperature products. In this paper, the combustion process will be investigated through the axis of a bi-propellant thruster by an in-house code with chemical reaction mechanism. This code includes several models for injection, droplet evaporation, liquid film, combustion and heat transfer through liquid and gas films. The Astrium thruster with MMH as fuel and NTO as oxidizer has been simulated. By implementing a detail mechanism with 1619 steps, the thruster has been simulated at different total mass flow rates and results have been validated by experimental data. Then, injector dimension effects on the droplet evaporation and combustion have been investigated. Results show that by increasing the injector dimension, the droplet evaporation length increases, so the flame structure changes in the combustion chamber. Therefore, the combustion products enter the nozzle with higher temperature and as a result, the thruster specific impulse increases.

کلیدواژه‌ها [English]

  • Thruster
  • Hypergolic
  • Swirl Injector
  • Monomethylhydrazine
  • Nitrogen Tetroxide

مراجع

  1.  

    1. A. Kakami, A. Kuranaga and Y. Yano, "Premixing-type liquefied gas bipropellant thruster using nitrous oxide/dimethyl ether," Aerospace Science and Technology, 94, 2019, 105351.
    2. J. R. Mason and R. D. Southwick, "Large liquid rocket engine transient performance simulation  system," Marshall Space Flight Center, Alabama, NASA CR-183780, 1989.
    3. J. Bradford, A. Charania and B. S. Germain, "REDTOP-2: Rocket Engine Design Tool Featuring Engine Performance, Weight, Cost, and Reliability," 40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit , Fort Lauderdale, Florida, AIAA-2004-3514, 2004.
    4. McBRIDE and GORDON, "Computer Program for Calculation of Complex Chemical Equilibrium Compositions, Rocket Performance, Incident and R efleaed Shocks, and Chapman-Jouguet Detonations," NASA SP-273, 1976.
    5. A. Ponomarenko, "RPA: Tool for Rocket Propulsion Analysis, Thermal Analysis of Thrust Chambers," software report, 2012.
    6. http://sierraengineering.com/ROCCID/roccid.html, Accessed 20 July 2020.
    7. K. J. Davidian, "Comparison of Two Procedures for Predicting Rocket Engine Nozzle Performance," in 23rd Joint Propulsion Conference, San Diego, CA, USA, AIAA-87-2071, 1987.
    8. C. Manfletti, "Start-Up Transient Simulation of a Pressure Fed LOx/LH2 Upper Stage Engine Using the Lumped Parameter-based MOLIERE Code," 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Nashville, 2010.
    9. H. L. Gray, "Modelling of combustion processes in small liquid bipropellant thruster," in 28th Joint Propulsion Conference and Exhibit, Nashville, 1992.
    10. M. EidiAttarZade, A. SarAbadani, G. Davarnia, H. Khosrobeygi, M. Farshchi and A. Ramezani, "Investigation of a Bi-propellant Thruster by a Developed Space Engine’s Thrust Chamber Analysis Code," Accepted in Journal of Space Science & Technology, 2020. )In persian(
    11. J. Hayashi, H. Tani, N. Kanno, D. Sato, Y. Daimon, F. Akamatsu and J. Gabl, "Multilayer reaction zones of a counterflow flame of gaseous Nitrogen Tetroxide and a liquid Monomethylhydrazine pool," Combustion and Flame, 201, 2019, pp. 244–251.
    12. S. Nonnenmacher and M. Piesche, "Design of hollow cone pressure swirl nozzles to atomize Newtonian fluids," Chemical Engineering Science , 55, No. 19, 2000, pp. 4339-4348.
    13. N. K. Rizk and A. H. Lefebvre, "Internal flow characteristics of simplex swirl atomizers," Journal of propulsion and power, 1, No. 3, 1985, pp. 193-199.
    14. S. Kim, T. Khil, D. Kim and Y. Yoon, "Effect of geometric parameters on the liquid film thickness and air core formation in a swirl injector," Measurement Science and Technology, 20, No. 1, 2008, 015403.
    15. N. K. Rizk and A. H. Lefebvre, "Prediction of velocity coefficient and spray cone angle for simplex swirl atomizers," Proceedings of the 3rd International Conference on Liquid Atomization and Spray Systems, London, 1985.
    16. H. Moongeun, J. Jeon and S. Y. Lee, "Discharge coefficient of pressure-swirl atomizers with low nozzle opening coefficients," Journal of Propulsion and Power , 28, No. 1 , 2012, pp. 213-218,.
    17. A. R. Jones, "Design optimization of a large pressure-jet atomizer for power plant," Proc. 2nd ICLASS, Madison, Wis., 1982.
    18. P. Fu, L. Hou, Z. Ren, Z. Zhang, X. Mao and Y. Yu, "A droplet/wall impact model and simulation of a bipropellant rocket engine," Aerospace Science and Technology, 88, 2019, pp. 32-39.
    19. H. Kang, H. Kim, S. Heo, S. Jung and S. Kwon, "Experimental analysis of hydrogen peroxide film-cooling method for nontoxic hypergolic thruster," Aerospace Science and Technology, 71, 2017, pp. 751–762.
    20. H. Tani, H. Terashima, Y. Daimon, M. Koshi and R. Kurose, "A Numerical Study on Hypergolic Combustion of Hydrazine Sprays in Nitrogen Tetroxide Streams," Combustion Science and Technology, 190, 2017, pp. 515–533.
    21. D. Preclik, O. Knab, D. Estublier and D. Wennerberg, "Simulation and Analysis of Thrust Chamber Flowfields: Storable Propellant Rockets," M. Popp, J. Hulka, V. Yang and M. Habiballah, Liquid Rocket Thrust Chambers, Virginia, AIAA, 2012, pp. 493-525.
    22. Z. Lian-bo, C. Min and X. Xu, "Performance Prediction of Apogee Attitude and Orbit Control Thruster for MMH/NTO Hypergolic Bipropellant," 50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Cleveland, OH, AIAA2014-3572.
    23. C. A. Salvador and F. S. Costa, "Vaporization Lengths of Hydrazine Fuels Burning with NTO," journal of propulsion and power, 22, 2006, pp. 1362-1372.
    24. K. H. Lee, "Numerical simulation on thermal and mass diffusion of MMH-NTO bipropellant thruster plume flow using global kinetic reaction model," Aerospace Science and Technology, 93, 2019, 104882.
    25. U. Gotzig and E. Dargies, "Development Status of Astriums New 22N Bipropellant Thruster Family," in 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, Alabama, 20-23 July 2003.
    26. D. Preclik, D. Estublier and D. Wennerberg, "An Eulerian-Lagrangian Approach to Spray Combustion Modeling for Liquid Bi-Propellant Rocket Motors," 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibi, San Diego, CA, July 10-12,1995.
    27. A. H. Lefebvre and V. G. McDonell, Atomization and Sprays, Taylor & Francis Group, Boca Raton, 2017.
    28. W. M. Grisson, Liquid Film Cooling in Rocket Engines, United states air force, Atlanta, Georgia, 1991.
    29. R. C. Stechman, J. Oberstone and J. C. Howell, "Film cooling design criteria for small rocket engines," 4th Propulsion Joint Specialist Conference, Cleveland, OH,1968.
    30. G. P. Sutton and O. Biblarz, Liquid, 7th ed., New York, John Wiley & Sons, pp. 197-240, 2001.
    31. J. D. Anderson, Modern compressible flow: with historical perspective, Boston, McGraw-Hill, 2003.
    32. M. R. Soltani, K. Ghorbanian, M. Ashjaee and M. R. Morad, "Spray characteristics of a liquid-liquid coaxial swirl atomizer at different mass flow rates," Aerospace science and technology, 9, No. 7, 2005, pp. 592-604.
    33. N. J. Labbe, Determining Detailed Reaction Kinetics for Nitrogen-and Oxygen-Containing Fuels, PhD Thesis, Department of Chemical Engineering, University of Massachusetts-Amherst, 2013.

     

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سابقه مقاله

  • تاریخ دریافت: 27 شهریور 1399
  • تاریخ بازنگری: 11 آبان 1399
  • تاریخ پذیرش: 01 دی 1399

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