DAQ Destroyer Hybrid Rocket Student Team: Travis Edwards, Viggo Hansen, Tom Slais, Clayton Chu, Michal Hughes, Guoshi Li, Gerard Finnegan, Tim Ip, Andrew Hatt, Ben Degang Faculty Advisors: Advisors: Carl Knowlen, Adam Adam Bruckner Bruckner,, James Hermanson, Tom Mattick University of Washington 7th Intercollegiate Rocket Engineering Competition Competition,, Green River,, UT, River UT, (June (Ju ne 21-23, 21-2 3, 2012) 201 2) Experimental Sounding Rocket Association, Logan, UT ESRA has permission to distribute
Introduction • Continuation of AA598/AA498 sounding rocket course • Program goal to develop a paraffin-N2O hybrid rocket motor
Introduction • Continuation of AA598/AA498 sounding rocket course • Program goal to develop a paraffin-N2O hybrid rocket motor
Design and Rationale • Minimum diameter • Modular upper airframe and fin can connect to oxidizer tank and combustion chamber • Scaled with estimated propellant load required to reach 25k ft AGL
Design and Rationale • Custom Paraffin-N2O hybrid motor • Inspired by recent advancements in hybrid propulsion • Rocket design scaled based on estimated motor performance • Payloads – Iridium spot II satellite locator – Downward facing GoPro camera
Hybrid Fuel Regression Rates • Traditional hybrid fuels(HTPB) have low regression rates, solid fuel is pyrolyzed directly to a gaseous phase limiting heat transfer • Paraffin fuel droplets are directly entrained from a liquid layer = Regression rate [mm/s] ṙ Oxidizer mass flux [kg/m2 – sec] = Regression rate coefficient [nona dimensional] n = Oxidizer mass flux exponent [nondimensional] = G o x
a
n
Paraffin-N2O Karabeyoglu [1]
0.155
0.5
50P-N2O T.S. Lee and H.L. Tsai [2]
0.1146
0.5036
HTPB-N2O Lohner et al. [3]
0.104
0.352
Systems Design, Analysis, and Testing
Liquid N2O Injector • 17-4PH Stainless Steel (62 mm O.D. x 6.4 mm thick) • Straight hole reamed orifices (1115 holes x 2 mm dia.) • 1.5kg/s LN2O mass flow • Coefficient of discharge 0.85
Ignition System • 2.4 mm thick polycarbonate diaphragm (burst P > 6.9 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
Avionics • Entirely student-designed and built avionics package, powered by Arduino • Array of sensors and communications • Redundant avionics and recovery system • Wireless connectivity between main avionics bay and combustion chamber
Sensor Package • Accelerometer – detects launch and coast • Magnetometer – detects rocket tipping (apogee) • Gyroscope • Barometric Pressure – detects altitude for main chute deployment • Oxidizer Tank Pressure • Combustion Chamber Pressure • GPS
Communications Package • Xbee-XSC 900 MHz – 14mi range – Communication to ground station
• Xbee-Pro 2.4GHz – short range – Communication to combustion chamber Arduino
• SIM Card – Send texts messages of detected events; GPS location
• UART – Access panel allows USB connection to rocket
Recovery • Dual Rouse Tech CD3’s deploy drogue • Main chute retained in deployment bag by tether • Tether released at appropriate altitude, deploying the main
Major Tests and Results
Propulsion • 7 full scale motor test firings • Specialized diagnostics employed – Gas Chromatography of motor exhaust – Flash X-ray of combustion chamber – Spectral analysis of combustion instabilities
• Major development issues: – Fuel grain integrity/composition – Fuel manufacturing/case bonding – Combustion chamber scaling (L*) – Chamber thermal protection
Full Scale Test Summary Test #
Fuel composition
1
Paraffin/ CB/HTPB, pour cast
2
Paraffin/ CB/HTPB, pour cast
3
Design/Test changes
Result
Notes
Combustor explosion
Igniter detonated
Reduced igniter mass
Low thrust
Suspected sloughing
Paraffin/ CB/HTPB, pour cast
Nozzle throat dia reduced
Low thrust
“”
4
Paraffin/ CB/HTPB, pour cast
None, demo at 6th IREC
Low thrust
“”
5
Paraffin/ CB/HTPB/ AL, pour cast
Injector pressure drop reduced, X-ray, ablative liners installed
Low thrust
Fuel grain failure mid burn
6
Paraffin/ CB/EAP, spun cast (Cracked grain)
L* increased, bell nozzle implemented, GC installed, DAQ system replaced
7
Paraffin/ Tar, spun cast (No cracks)
DAQ system replaced again
DAQ failure after 35 ms Design thrust reached
DAQ failure after 1.9 s
Test #3 • Abrupt change in thrust mid burn • Sloughing and or flame holding instability suspected
Test #5 • 150 kV flash X-ray system install • Combustor design modified – Increased fuel loading – Modified pre-combustor – Injector area increased – -Ablative phenolic liners installed
Test # 5 Flash X-Ray averaged regression rates • 2.3 mm/s preburn vs. postburn • With simulated oxidizer mass flux model and Stanford regression rate coefficients(a = 0.155 n = 0.5) averaged regression rate is 2.7 mm/s
Fuel Development Revisited • Sensitive manufacturing process – Spin casting appears to be required
• Additives required for structural integrity: – HTPB, Ethylene-vinyl acetate, Vybar 103, Tar
• IR opaque additives used to increase heat transfer to fuel surface: – Carbon black, Tar
Fuel Development Revisited • Pour casting of paraffin/HTBP mixtures possible with advanced process control [2] • Spin casting of paraffin with HTPB causes stratification • Carbon black will stratify in paraffin at > 420 rpm • Tar remains mixed with paraffin at higher rpm (820)
Test #6 • Fuel grain spun cast with Ethylene-vinyl acetate and tar • Fuel grain cracked during mfg but remained case bonded • Gas Chromatography Sample station installed • Combustor L* increased from 4.3 to 6.8 m • Bell nozzle designed with method of characteristics installed
Characteristic Combustor Length L* • L* is the ratio of average combustor volume to nozzle throat area. • L* is related to residence time in the combustion chamber. • Without a sufficiently large L* there will be incomplete combustion. • Hybrid motors require ~10 times the L* of a liquid bipropellant rocket for acceptable combustion efficiency [1]
Lengthened combustion chamber • High regression rate fuel favors short fuel grain • Increased L* from lengthened post combustor
L* = 4.3m
L* = 6.8 m
Test #6 • DAQ system crashed 35 ms after ignition • Initial data showed high performance • Flash X-ray failed to trigger • Gas chromatography sample station was triggered at an unknown time
Gas Chromatography Sample • Analyzed with flame ionization and thermal conductivity detectors • CO/CO2 ratio allows estimate of O/F ratio via correlation with chemical equilibrium(NASA CEA) simulation
Test #6 chromatogram
Test #6 GC results • 16% CO, 7% CO2, and 11% H2 detected • Detected CO/CO2 ratio corresponds to an O/F of 6.7 according to CEA • Close to optimum O/F ratio for nitrous oxide paraffin of 7
Test #7 • DAQ system replaced • Fuel grain cracks resolved with 10% tar additive • All other test conditions held constant
Test #7
Test #7 results • Laptop free fall protection disabled the hard drive after 1.9 seconds • Target chamber pressure and thrust reached (400 psi, 800 lbf) • Combustion instability • Design thrust reached • Flash X-ray triggered as planned 3 seconds after ignition • GC sample failed to trigger
Test # 7 results • Cracks observed in fuel grain after ignition however remained case bonded • Fuel composition regressed at a rate 105% predicted with Stanford regression rate coefficients and calculated oxidizer mass flow rate • Instantaneous ISP estimated 220-230 seconds
Test #7 Combustion Instability • Significant combustion instability observed • Spectral analysis of load cell data showed the first two longitudinal acoustic modes • Low frequency modes consistent with intrinsic low frequency instabilities of the hybrid combustion process and/or feed system coupled instabilities
Final Design Summary • Hybrid propulsion system validated • Initial thrust sufficient to safely leave the launch rail with a stability margin of 2 calibers • Open source sensor suite and camera included to observe propulsion system
Acknowledgements
• Special thanks: – Aerojet/Gencorp Foundation – David Stechmann – Aero-Astro, Chemistry, Physics and Mechanical Engineering Machine Shops
