《航天推进 Space Propulsion》（英文版）Lecture 9: Some Examples of Small Solid

16.522, Space Propulsion Prof. Manuel martinez-Sanchez Lecture 9: Some Examples of Small Solid Propellant Rockets for In-space Propulsion The sTAR 13B incorporates the lightweight case developed for the staR 13 with the propellant and nozzle design of the earlier TE-M-516 apogee motor. The motor case has been stretched 2. 2 inches to provide for increased propellant loading the motor has been used to adjust orbit inclination of a satellite from a Delta laun MOTOR PERFORMANCE (70 F Vacuum) Burn Time/Action Time, sec 148/161 Ignition Delay Time, sec 0.02 Burn Time Average Chamber Pressure, psia 823 Action Time Average Chamber Pressure, psia 787 Maximum Chamber Pressure, psia 935 Total Impulse, Ibf-sec 26,040 Propellant: Specific Impulse, Ibf-sec/lbl 286.6 Effective Specific Impulse, Ibf-sec/Ibm Burn Time Average Thrust, Ibf 1708 Action Time Average thrust, ibf 1577 Maximum thrust ibf 2160 SPIN CAPABILITY, rpm 120 CASE 6A1-4V Titanium Minimum Ultimate Strength, psi 165,000 Minimum Yield Strength, psi 152,000 Hydrostatic Test Pressure, psi 1330 Yield pressure, psi Hydrostatic Test Pressure/Maximum Pressure 1.05 Nominal thickness in 0.035 NOZZLE Exit Cone material Vitreous silica phenolic Throat insert material ATJ Graphite Initial Throat Area, in2 Exit diameter in 8.02 Expansion Ratio, Initial/Average 49.8/41 Expansion Cone Half Angle, deg TyI Fixed Number of nozzles LINER Density Ibm/in 0.045 16.522, Space P pessan Lecture 9 Prof. Manuel martinez Page 1 of 3

16.522, Space Propulsion Prof. Manuel Martinez-Sanchez Lecture 9: Some Examples of Small Solid Propellant Rockets for In-space Propulsion The STAR 13B incorporates the lightweight case developed for the STAR 13 with the propellant and nozzle design of the earlier TE-M-516 apogee motor. The motor case has been stretched 2.2 inches to provide for increased propellant loading. The motor has been used to adjust orbit inclination of a satellite from a Delta launch. MOTOR PERFORMANCE (70 ° F Vacuum) Burn Time/Action Time, sec 14.8/16.1 Ignition Delay Time, sec 0.02 Burn Time Average Chamber Pressure, psia 823 Action Time Average Chamber Pressure, psia 787 Maximum Chamber Pressure, psia 935 Total Impulse, lbf-sec 26,040 Propellant: Specific Impulse, lbf-sec/lbm 286.6 Effective Specific Impulse, lbf-sec/lbm 285.7 Burn Time Average Thrust, lbf 1708 Action Time Average Thrust, lbf 1577 Maximum Thrust, lbf 2160 SPIN CAPABILITY, rpm 120 CASE Material 6Al-4V Titanium Minimum Ultimate Strength, psi 165,000 Minimum Yield Strength, psi 152,000 Hydrostatic Test Pressure, psi 1330 Yield Pressure, psi 1394 Hydrostatic Test Pressure/Maximum Pressure 1.05 Nominal Thickness, In. 0.035 NOZZLE Exit Cone Material Vitreous Silica Phenolic Throat Insert Material ATJ Graphite Initial Throat Area, in2 1.14 Exit Diameter, In. 8.02 Expansion Ratio, Initial/Average 49.8/41.0 Expansion Cone Half Angle, deg 17 Type Fixed Number of Nozzles 1 LINER Type TL-H-304 Density, lbm/ in. 3 0.045 16.522, Space Propulsion Lecture 9 Prof. Manuel Martinez-Sanchez Page 1 of 3

IGNITION TRAIN Components S&A/etA/TBi/pyrogen igniter Minimum Firing Current per Detonator, amperes 5.0 Circuit Resistance per Detonator, ohms 1.0 No, of detonators and tbis Squib or Tbi compatible WEIGHTS Ibm Total Loaded 103.7 Propellant 90.9 Case Assembly 5.64 Nozzle assembl Igniter Assembl 0.68 Internal Insulation 2.34 Liner 0.14 Miscellaneous 0.28 Total Inert(excluding igniter propellant) 12.80 Burnout 12.30 Propellant Mass Fraction 0.87 TEMPERATURE LIMITS Operation 40to+110°F Storage 40to+110°F PROPELLANT Propellant Designation and Formula TP-H-3082 AP-70 A|-16% CTPB Binder-14%/ PROPELLANT CONFIGURATION Type Internal burning 8-point star Web. it Web fraction %/o Silver fraction %o Propellant Volume, in3 1446 Volumetric Loading Density 92 Web Average Burning surface Area, in 345 Initial Surface to throat Area ratio 316 PROPELLANT CHARACTERISTICS Burn Rate at 1000 psia, in /sec 0.301 Burn rate Exponent 0.31 Density Ibm/i 0.0628 Temperature Coefficient of Pressure, %o/F Characteristic Exhaust Velocity, ft/sec 5025 Adiabatic Flame Temperature, F 16.522, Space Propulsion Lecture 9 Prof. Manuel martinez-Sanchez Page 2 of 3

IGNITION TRAIN Components Minimum Firing Current per Detonator, amperes Circuit Resistance per Detonator, ohms No. of Detonators and TBIs Squib or TBI compatible S&A/ETA/TBI/pyrogen igniter 5.0 1.0 2 WEIGHTS, lbm Total Loaded 103.7 Propellant 90.9 Case Assembly 5.64 Nozzle Assembly 3.72 Igniter Assembly 0.68 Internal Insulation 2.34 Liner 0.14 Miscellaneous 0.28 Total Inert (excluding igniter propellant) 12.80 Burnout 12.30 Propellant Mass Fraction 0.87 TEMPERATURE LIMITS Operation 40 to +110°F Storage 40 to +110°F PROPELLANT Propellant Designation and Formula TP-H-3082 AP-70% Al-16% CTPB Binder-14% PROPELLANT CONFIGURATION Type Web, In. Web Fraction, % Silver Fraction, % Propellant Volume, in.3 Volumetric Loading Density Web Average Burning Surface Area, in.2 Initial Surface to Throat Area Ratio Internal burning, 8-point star 4.187 62 2 1446 92 345 316 PROPELLANT CHARACTERISTICS Burn Rate at 1000 psia, in./sec 0.301 Burn rate Exponent 0.31 Density, lbm/in.3 0.0628 ° Temperature Coefficient of Pressure, %/ F 0.10 Characteristic Exhaust Velocity, ft/sec 5025 ° Adiabatic Flame Temperature, F 5662 16.522, Space Propulsion Lecture 9 Prof. Manuel Martinez-Sanchez Page 2 of 3

Effective Ratio of Specific Heats(chamber) 1.16 (NoZzle Exit) 1.21 CURRENT STATUS Production BC1355B4/91 2000 TIME sec Star 13 B TE-M-763 148-KS-1,708 Orbit Insertion Motor 13.57 DIA 802 25.11 16.522, Space P pessan Lecture 9 Prof. Manuel martinez Page 3 of 3

Effective Ratio of Specific Heats (chamber) 1.16 (Nozzle Exit) 1.21 CURRENT STATUS Production BC1355B 4/91 16.522, Space Propulsion Lecture 9 Prof. Manuel Martinez-Sanchez Page 3 of 3