Solid Rocket Motor Design for an Undergraduate Built

Solid Rocket Motor Design for an Undergraduate Built Sounding Rocket C 1 C Richard Tuminello and Major Dan Miller U. S. Air Force Academy, Department of Astronautics

Abstract The U. S. Air Force Academy rocket design course has been developing a 100 -lb rocket to go to a 100 -km altitude with a 5 -lb payload of scientific instruments Project Falcon. Launch is currently building a sub-performance vehicle to launch to an altitude of 30, 000 feet

Teams 3 Teams of several students Payload l Aero Vehicle l Propulsion l l The thrust chamber is the focus of this presentation

Introduction This will focus on the Propulsion Team’s Design l Tests l 100 lb thrust test l Full-scale motor test l Prototype rocket launch l

Propellant Grain Propellant Weight: 20 lbs Diameter: 5 inches Port Diameter: 2 inches

Propellant Shape “Moon Burn” Bore is tangent to case Regressive Burn with over 700 lbs thrust initially Gives greater than 5 to 1 thrust to weight ratio to ensure fin stabilization off the launch rail

Burn Rate: 0. 368 inches/second l Achieved the desired mass flow rate and thrust profile of the rocket.

Insulation ¼ inch thick Cured EPDM rubber l Non-fiber reinforced Chemlock adhesive used to bind insulation to inside of case

Thrust Chamber Case Aluminum 6061 -T 6 Outer diameter: 6 inches Inner diameter: 5. 5 inches Wall thickness: 0. 25 inches

Thrust Chamber End Caps Aluminum 6061 -T 6 Vitton rubber o-rings to hold pressure Glass phenolic insulation Steel role pins used to secure end caps to the case Ablative RTV used to seal rubber to phenolic inside case

Rocket Motor Assembly

Nozzle Chopped fiberglass phenolic One solid piece design 15 degree half angle

Ignition Method Piccolo tube screws into north end cap Pellets of Boron Potassium Nitrate Squib initiated remote ignition l Shoots a flame down the port

100 lb Thrust Motor Test Used primarily to obtain feasibility of the rubber insulation l Regression rate of EPDM rubber is 0. 00112 in/sec

Full-Scale Motor Test Objectives: Analyze thrust profile l Test the piccolo tube ignition method obtain l Test the EPDM regression rate l Obtain the phenolic regression rates l Test the reliability of the nozzle assembly l

Thrust Profile Expected total burn time: 10. 4 sec Greater than 500 lbs for first 4. 26 seconds to ensure fin stabilization off the launch rail Spike at about 3. 36 sec is where throat insert was expelled out the south end

Piccolo Tube Survived with only charring No burning past the glass phenolic insulation it was screwed into

Rubber Insulation EPDM rubber after fullscale test fire Gas burned through by the south end because of higher flow rates and possible air pockets where the insulation did not bond completely to the case

Chamber Pressure Maximum pressure is 1275 psi Pressure dropped off drastically because of the failure of the throat insert in the nozzle

Important Equations Initial mass flow rate l 1. 56371 kg/sec Characteristic exhaust velocity (c*) l l Theoretical=1537 m/sec Experimental=1397 m/sec Specific Impulse (Isp) l l Theoretical=254 sec Experimental=218 sec

Conclusions and Recommendations Problems present during the testing phase have been addressed and should no longer be an issue Rubber insulation l Nozzle material l Ignition method l Project Falcon. Launch is a “GO” on 6 April 2003 to an altitude of 30, 000 feet

Questions?