1 / 34

Development of an O-Class Paraffin/HTPB-N 2 O Hybrid Rocket Motor

Development of an O-Class Paraffin/HTPB-N 2 O Hybrid Rocket Motor. Propulsion Team V. Hansen T. Edwards M. Hughes Advisors A.P. Bruckner J. Hermanson C. Knowlen A.T. Mattick Sponsored By GenCorp. Presented at the 2011 AIAA Pacific Northwest Technical Symposium.

hyman
Download Presentation

Development of an O-Class Paraffin/HTPB-N 2 O Hybrid Rocket Motor

An Image/Link below is provided (as is) to download presentation Download Policy: Content on the Website is provided to you AS IS for your information and personal use and may not be sold / licensed / shared on other websites without getting consent from its author. Content is provided to you AS IS for your information and personal use only. Download presentation by click this link. While downloading, if for some reason you are not able to download a presentation, the publisher may have deleted the file from their server. During download, if you can't get a presentation, the file might be deleted by the publisher.

E N D

Presentation Transcript

  1. Development of an O-Class Paraffin/HTPB-N2O Hybrid Rocket Motor Propulsion Team V. Hansen T. Edwards M. Hughes Advisors A.P. Bruckner J. Hermanson C. Knowlen A.T. Mattick Sponsored By GenCorp Presented at the 2011 AIAA Pacific Northwest Technical Symposium

  2. Introduction Ongoing development effort for: Experimental Sounding Rocket Association (ESRA) Intercollegiate Rocket Engineering Competition (IREC) Held annually in Green River Utah http://www.soundingrocket.org/

  3. UW Hybrid Rocket Project • 1 quarter design, 1 quarter manufacturing and testing • 498/598 Special topics class • Project Goal: Enter a paraffin hybrid rocket in the advanced class of the ESRA competition • Advanced Class Competition Goal: Launch to 25,000 ft AGL with 10lb payload, recover • Minimum O-class motor required (20-40 kN-s)

  4. Why use Paraffin? • HTPB based hybrids have low regression rates • multi-port fuel grains required, poor volumetric efficiency • Recent advancement’s in hybrid propulsion at Stanford using liquefying fuels (liquid layer theory) • Order of magnitude increases in regression rate • Allows single port fuel grains • Hybrids have environmental, safety and simplicity advantages • Paraffin hybrids are competitive with RP1-LOX for orbital spaceflight

  5. Paraffin Caveats • Sensitive manufacturing process • 9-12% Shrinkage makes form casting impractical, requires spin casting • Brittle Structural can result in sloughing • Structural additives necessary e.g Low Density Polyethylene Wax (Vybar 103), HTPB • Carbon Black addition as an optical opacifier to increase heat transfer to the liquid layer

  6. Development Plan • Subscale regression rate study to test Paraffin fuels with various additives • Schedule requires rocket & motor to be scaled and begin manufacturing in parallel with subscale tests • Verify motor performance with full scale static fires • Systems integration of the propulsion system

  7. Subscale Testing Results • Total of 27 subscale static fires w/ N2O & GOx • Fuel compositions tested with N2O • 92% paraffin, 6% LDPE wax (Vybar 103), 2% carbon black (test matrix of 11 tests) • 62% paraffin, 30% 3 m aluminum , 6% LDPE wax, 2% CB (2 tests) • 50% paraffin, 50% HTPB (50P) (1 test)

  8. Subscale Testing Results • Oxidizer mass flow calibrated across custom flow restrictor (Venturi/ANSI orifice was not used) • Space and time averaged N2O-paraffin regression rate measured and compared with published data • Outliers caused by procedural errors removed from curve fit ṙ = Regression rate [mm/s] Gox= Oxidizer mass flux [kg/m2 – sec] a = Regression rate coefficient [non-dimensional] n = Mass flux exponent [non-dimensional]

  9. N2O Tank Testing • Water flow tests used to determine injector Coefficient of discharge (0.8) • Hydrostatic testing • First weld attempt failed during hydro-test due to insufficient weld penetration • Second hydro-test with revised weld geometry reached 9.65 MPa Water flow injector test

  10. Liquid N2O Injector • 17-4PH Stainless Steel (62 mm O.D. x 6.4 mm thick) • Straight hole reamed orifices (11 or 13 holes x 2 mm dia.) • 1.5kg/s LN2O mass flow • Coefficient of discharge 0.8

  11. Ignition System • 2.4 mm thick polycarbonate diaphragm (burst P > 6.894 MPa) • Custom pyrotechnic fixed to hydrostatically pre-domed diaphragm • KClO4 (60%) + GEII Silicone (20%) + 3 m Al (20%) • Experimental regression rate: 25 mm/s • Pyrotechnic ruptures diaphragm and ignites rocket motor

  12. Ignition System Test w/CO2

  13. Motor Testing • Scaling issues with spin casting full scale grain arose just prior to first test • Addition of 10% HTPB allowed form casting • Styrofoam mandrel dissolved with acetone post-casting • 88% Paraffin, 10% HTPB, 2% Carbon black

  14. Motor Testing (cont) • 4 full-scale static tests to date • Test #1 • Sustainer in pre-combustor burned faster than expected/ possibly detonated • Combustion chamber over pressurized to failure

  15. Full-scale Motor Test #1

  16. Full-scale Motor Test #1 (cont)

  17. Motor Testing (cont) • Combustor rebuilt with spare parts in 2 days • Sustainer removed on later tests • Pyrotechnic on diaphragm by itself proved to be sufficient for motor ignition • Test #2 & #3 • Reliability/repeatability verified • Combustor parameters, fuel load, unchanged • High O/F ratio ~10 • Test #4 Demo at 6th IREC • Implemented CO2 purge

  18. Full-scale Motor Test #2 Video

  19. Full-scale Motor Test #4 Video

  20. Motor Testing (cont) • Several factors believed to have cause lower than design thrust • High O/F ratio of ~10 • Flame holding Instability, possibly due to lack of step on the front of fuel grain • Onset of stability coincides with front of the fuel grain burning through to the combustor wall

  21. Combustion Chamber (as tested) • Phenolic liners at pre- and post-combustor ends • Cotronics alumina ceramic adhesive layer inside cardboard liner • Graphite nozzle (4:1 expansion ratio)

  22. Future work • Full-scale test program • X-ray diagnostic for regression rate measurement • Test Grade C Phenolic motor liner for improved thermal protection • Increase fuel load, redesign combustor layout • Test aluminized fuel • Impinging injector • Achieve stable combustion • Subscale motor test program • Test structurally robust grain compositions • High accuracy calibrated Venturi oxidizer flow measurement

  23. Questions ? References: [1] Karabeyoglu, M.A., Zilliac, G., Castellucci, P., Urbanczyk, P., Stevens, J., Inalhan, G., and Cantwell, B.J. “Development of High-Burning-Rate Hybrid-Rocket-Fuel Flight Demonstrators” 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Huntsville, AL, July 2003. [2] T.S. Lee and H.L. Tsai, “Fuel Regression Rate in a Paraffin-HTPB Nitrous Oxide Hybrid Rocket,” 7th Asia-Pacific Conference on Combustion, National Taiwan University, Taipei, Taiwan, 24-27 May, 2009 [3] Kevin Lohner et al., “Fuel Regression Rate Characterization Using a Laboratory Scale Nitrous Oxide Hybrid Propulsion System,” AIAA-2006-4671, 2006. Special thanks to the Aeronautics & Astronautics , Mechanical Engineering, Physics and Chemistry machine shops and Automobili Lamborghini ACSL Contact Email: vkhansen@gmail.com Team Website: www.sarpuw.org

  24. Backup Slides

  25. Fuel Compositions Plotted • Karabeyoglu - Paraffin, 1 % Carbon black and proprietary structural additives, spun cast • T.S. Lee and H.L. Tsai - 50 % Paraffin, 46 % HTPB, 4 % IDPI, progressively poured and cooled then annealed for 7 days at 40C, form cast • Experimental test matrix - 92 % Paraffin, 6 % Vybar 103 (LDPE wax), 2 % Carbon black, spun cast • Experimental test 50P – 50 % Paraffin, 46 % HTPB, 4 % IDPI , form cast

  26. Ignition System Testing • Diaphragm with pyrotechnic • Determine optimal pyrotechnic composition and placement • Pre-domed diaphragm does not separate from pyrotechnic • Pyrotechnic ignites motor after diaphragm bursts

More Related