AAE STAGING. Ch3 46

Transcription

1 3.7 STAGING Ch3 46

2 STAGING PHILOSOPHY Staging Strategy: Reduces effective mission-averaged mass ratio. Serial staging is more effective. Parallel staging ( strap-ons ) simpler, especially to upgrade existing system. # of stages must be balanced with additional complexity. Staging Configurations Definitions: Booster Stage First (lowest) stage, Largest stage providing the largest thrust and total impulse. Sustainer Stage Upper (subsequent) stage, Thrust becomes smaller. Half Stage or Zero Stage Boosters operate in parallel to first stage, Strap-on solid propellant motors. Ch3 47

3 NOMENCLATURE Vehicle consists of N stages, Stage: j = any of the stages, Mass Definitions for arbitrary stage n < N: m 0n = initial mass of stage n, m fn = burnout mass of stage n, m PL = payload mass of the last stage to fire, 1 j N Structural Ratio (Coefficient, Factor): ε j = m f j m 0 j NOTE for top stage where n=n ε N = m fn m PL m 0N m PL Payload Ratio (Coefficient, Factor): NOTE for top stage where n=n λ N = m PL m 0N λ j = N m 0 j j+1 N m 0 j j Ch3 48

4 DEFINITIONS Mass Ratio Definitions for Staged Vehicle: Stage Mass Ratio R (MR): Mass of the vehicle at ignition of the j th stage divided by its mass at burnout of the j th stage N 1 m R i=j 0i j = R j = ε j (1 λ j ) + λ j N m f j + m i=j+1 0i MR j = ε j (1 λ j ) + λ j Propellant Mass Fraction is a measure of the structural efficiency, or what proportion of the mass of the system goes toward developing thrust and moving a payload. Stage Velocity Increment: Δv j = u eq,j lnr j = g 0 I sp,j ln 1 MR j Final Velocity Increment Δv f = N j=1 Δv j ζ j = m Pj m 0 j = m Pj m Pj + m inert,j = Δv 1 + Δv 2 + Δv 3 + = u eq,1 lnr 1 + u eq,2 lnr 2 + u eq,3 lnr 3 + Ch3 49

5 STAGING PAYOFF Effect of Staging on Performance Ch3 50

6 EXAMPLE Ch3 51

7 Staging - Examples World-wide launcher scenario 2001 for payloads above 3000 kg in GTO. Launcher Delta III Atlas2AS H2A Proton AR44L CZ3B Zenit 3 Titan4B AR5 Delta IVH Country USA USA Japan Russia Europe China Ukraine USA Europe USA Ch3 52

8 3.8 VEHICLES Ch3 53

9 Launch Vehicle Ch3 54

10 Atlas Trajectory Profile Ch3 55

11 Launch Vehicle Saturn V Liftoff Thrust: 33, kn Total Mass: 3,038,500 kg Core Diameter: m Total Length: m Launches 13 Success Rate: 100% First Launch Date: 09 Nov 1967 Last Launch Date: 14 May 1973 LEO Payload: 118,000 kg to 185 km orbit at 28.0 degrees. Payload: 47,000 kg to a translunar trajectory. Development Cost: $7,439.60M in 1966 average dollars. Launch Price: $ 431M in 1967 dollars. Ch3 56

12 Launch Vehicle STAGES OF THE SATURN V BOOSTER FIRST STAGE SECOND STAGE THIRD STAGE APOLLO DIAMETER 33' 33' 21' 8" 21' 8" LENGTH 138' 81' 6" 58' 7" 81' 10¼" WEIGHT 4,792,000 lbf (fueled) 1,037,000 lbf (fueled) 262,000 lbf (fueled) 109,608 lbf ENGINES five F-1 rocket engines five J-2 rocket engines one J-2 rocket engine - PROPELLANT kerosene & LOX LH 2 & LOX LH 2 & LOX - THRUST 7.5 million lbf 1 million lbf 200,000 lbf - 1st BURNTIME 2 min, 50 sec 6 min, 35 sec 2 min, 45 sec - VELOCITY at 1st burnout 6,200 mph 15,400 mph 17,500 mph - ALTITUDE at 1st burnout 36 miles miles 115 miles - 2nd BURNTIME min, 12 sec - VELOCITY at 2nd burnout ,300 mph Translunar injection to 60 nmi above moon - ALTITUDE at 2nd burnout Ch3 57

13 Saturn V - Stage 1 Gross Mass: 2,286,217 kg Empty Mass: 135,218 kg Thrust (vac): 38,716 kn Isp: 304 s (vac), 265 s (SL) Burn time: 161 sec Propellants: Lox/Kerosene Diameter: m Span: 19.0 m Length: 42.1 m No. Engines: 5, F-1 Thrust (vac): 7,7433 kn Mass Engine: 8,391 kg Diameter: 3.72m Length: 5.64m Chambers: 1 Chamber Pressure: 70 bar Area Ratio: Oxidizer to Fuel Ratio: 2.27 Thrust to Weight Ratio: Ch3 58

14 Saturn V - Stage 2 Saturn II Gross Mass: 490,778 kg Empty Mass: 39,048 kg Thrust (vac): 5,1676 kn Isp: 421 s (vac), 200 s (SL) Burn time: 390 s Propellants: Lox/LH2 Diameter: m Span: m Length: m No. Engines: 5, J-2 Thrust (vac): 1,033 kn Burn Time: 475 s Mass Engine: 1,438 kg Diameter: 2.01 m Length: 3.38 m Chambers: 1 Chamber Pressure: 30 bar Area Ratio: 28 Oxidizer to Fuel Ratio: 5.50 Thrust to Weight Ratio: Ch3 59

15 Saturn IVB Saturn V - Stage 3 Gross Mass: 119,900 kg Empty Mass: 13,300 kg Thrust (vac): 1,032 kn Isp: 421 s (vac), 200 s (SL) Burn time: 475 s Propellants: Lox/LH2 Diameter: 6.61 m Span: 6.61 m Length: 17.8 m No. Engines: 1, J-2 Thrust (vac): 1,033 kn Mass Engine: 1,438 kg Diameter: 2.01 m Length: 3.38 m Chambers: 1 Chamber Pressure: 30 bar Area Ratio: 28 Oxidizer to Fuel Ratio: 5.50 Thrust to Weight Ratio: Ch3 60

16 Soviet Moon Rocket Stage 1 Stage 2 Stage 3 N A N B N V Gross Mass: 1,942,000 kg 506,000 kg 193,000 kg Empty Mass: 192,000 kg 50,000 kg 16,000 kg Thrust (vac): 49,442.4 kn 13,734 kn 1,569.6 kn Isp: 331 s, 296 s (SL) 346 s 347 s Burn time: 113 s 106 s 368 s Diameter: m 6.80 m 4.80 m Span: m 9.80 m 6.80 m Length: m m m Propellants: LOX/Kerosene LOX/Kerosene LOX/Kerosene No Engines: 30, NK-15 8, NK-15V 4, NK-19 Status: Design 1964 Design 1964 Design 1964 Ch3 61

17 Inertial Upper Stage (IUS) IUS is a 2-stage, inertially guided, three axis stabilized solid rocket: Size: 17 long and 9.25 in diameter, Mass: 32,500 lb. 1 st stage (SRM and Interstage): Burn time: approx. 2 min burn from LEO to GTO, Propellant mass: 21,580 lb, Average thrust: 4,000 lbf. 2 nd stage includes: equipment support section (GNC, telemetry, command and data management, reaction control and electrical power), Solid Rocket Motor (SRM) Propellant mass: 6,000 lb Average thrust: 18,200 lbf. Ch3 62

18 Inertial Upper Stage (IUS) p c = 4.21 MPa = 181 I sp = 304 s t b = 101 s p c = 4.52 MPa = 64 I sp = 296 s t b = 142 s Ch3 63

19 Centaur RL-10B Thrust (vac): 110 kn Isp: 462 s Burn Time: 700 s Diameter: 2.13 m Area Ratio: 250 Dual-engine upper stage Centaur for Delta, Atlas, Titan rockets. Powered by two Pratt & Whitney RL10A-4-2 turbopump-fed engines burning liquid oxygen and liquid hydrogen. GNC, tank pressurization, and propellant usage controls for both boost and Centaur phases are provided by the inertial navigation unit (INU) located on the forward equipment module. First Centaur burn lasts about five minutes, after which the Centaur and its payload coast in a parking orbit. During the first burn, approximately eight seconds after ignition, the payload fairing is jettisoned. Second Centaur ignition occurs approx. 30 minutes into the flight, continues for about three minutes, and is followed several minutes later by the separation of the spacecraft from Centaur. Ch3 64

20 Apogee Engines LEROS 1 Designed and manufactured by ARC UK, Dual mode apogee engines utilizing hydrazine and nitrogen tetroxide propellants at a thrust level of 100 lbf (445 N), Over 50 engines have been delivered since the first LEROS 1 engine was developed for the NEAR spacecraft, now orbiting asteroid (Eros 433), LEROS 1B Design improvements raised specific impulse from 315 to 320 s, LEROS 1C First flown in 1999, Specific impulse of 325 s, Highest performance apogee engine flown to date. Ch3 65

21 Mag-Lifter Launch assist system for a single-stage-to-orbit (SSTO) highly reusable vehicle (HRV): Magnetic levitation guideway runs up the side of a mountain and releases the vehicle at an altitude of 10,000 ft (3.0 km) to minimize air drag, Deceleration guideway at the end of the launch tube to bring the maglev accelerator-carrier to a stop for re-use, Final exit velocity can range from 600 to 1,000 mph (0.3 to 0.5 km/s), SSTO's engines can be started prior to release for checkout, Deceleration guideway is sized to bring the combined maglev accelerator-carrier/vehicle to a stop Figure courtesy Joe Howell for launch abort, NASA MSFC Launch tube can be partially evacuated (to minimize air drag in the tube) with a membrane used to seal the exit end of the tube. Ch3 66

22 Supergun - Project Babylon Developed for Iraqi military by Gerald Bull, presumed to be assassinated by Israeli agents in Nine tons of special supergun propellant could fire a 600 kg projectile over a range of 1,000 km, or a 2,000 kg rocket-assisted projectile. Barrel was 156 m long, with a 1 m bore. The 2,000 kg projectile would place a net payload of about 200 kg into orbit at a cost of $ 600 per kg. Ch3 67

23 Supergun - Project Babylon From March of 1988 until the invasion of Kuwait in 1990, Iraq contracted with Gerard Bull to build three superguns: two full sized 'Project Babylon' 1000 mm guns and one 'Baby Babylon' 350 mm prototype. Nine tonnes of special supergun propellant could fire a 600 kg projectile over a range of 1,000 kilometres, or a 2,000 kg rocket-assisted projectile. The 2,000 kg projectile would place a net payload of about 200 kg into orbit at a cost of $ 600 per kg. The 1000 mm guns were never completed. After the war UN teams destroyed the guns and gun components in Iraqi possession. LEO Payload: 200 kg. to: 180 km Orbit. at: 33.0 degrees. Payload: 600 kg. to a: 500 km altitude suborbital trajectory. Total Mass: 2,000 kg. Core Diameter: 1.00 m. Launch Price $: 0.12 million. in 1990 price dollars. Total Development Built: 25. Gerard Bull was first contracted by the Iraqi government in They desperately needed artillery to out-range the enemy in the protracted Iran -Iraq War. The association allowed Bull to yet again seek sponsorship of a space-launch supergun. Sadam Hussein liked the idea, and Project Babylon was born. Throughout the 1980's Bull's dealings with Iraq had the covert approval of Western governments, who saw Iraq as a counterweight to revolutionary Iran. In March of 1988, Bull received a contract to build two full sized 'Project Babylon' 1000 mm superguns and one 'Baby Babylon' 350 mm prototype for a total of $25 million. The project was given the cover designation 'PC-2' (Petrochemical Complex-2). British engineer Christopher Cowley was the project manager. The Project Babylon gun would have a barrel 156 meters long with a one meter bore. The launch tube would be 30 cm thick at the breech, tapering to 6.5 cm at the exit. Like the V-3 the gun would be built in segments. 26 six-meter-long sections would make up the barrel, totalling 1510 tonnes. Added to this would be four 220 tonne recoil cylinders, and the 165 tonne breech. The recoil force of the gun would be 27,000 tonnes - equivalent to a nuclear bomb and sufficient to register as a major seismic event all around the world. Nine tonnes of special supergun propellant would fire a 600 kg projectile over a range of 1,000 kilometres, or a 2,000 kg rocket-assisted projectile. The 2,000 kg projectile would place a net payload of about 200 kg into orbit at a cost of $ 600 per kg. In May of 1989 the Baby Babylon was completed at Jabal Hamrayn, 145 km north of Baghdad. The horizontally-mounted gun was 45-m long with a 350 mm barrel, and had a total mass of 102 tonnes. Following tests using lead projectiles the gun was reassembled on a hillside at a 45 degree angle. It was expected to achieve a range of 750 km. An Iraqi defector revealed later that the gun was to be used for several missions: Long-range attack using chemical, biological, or nuclear warheads. However since the weapon was fixed, it could only be fired in one direction, and like the V-3 would be easily identified and neutralised by the targeted country. For this reason the Israelis did not consider it much of a threat. As an anti-satellite weapon. It would launch a special shell in space that would explode near the target satellite, covering it with sticky material and blinding it. Production of components for the Babylon gun began in Britain with the knowledge of British and US intelligence services. These were officially 'oil pipeline segments' for the 'PC-2' petrochemical refinery. On 22 March 1990 Bull was assassinated and the project quickly unravelled. It is commonly thought that he was killed by the Israelis, concerned not so much by the supergun work but rather dynamics research Bull was doing to improve Iraqi ballistic missiles. Three weeks later British Customs seized the final eight sections of the Babylon Gun. On August 2, 1990, Iraq invaded Kuwait, precipitating the Gulf War. This ended Western covert sponsorship of Iraq. After the war UN teams destroyed the Baby Babylon 350 mm gun, components of the two Project Babylon 1000 mm guns (44 gun sections, four recoil cylinders), and one tonne of supergun propellant (the Iraqis destroyed the remaining 11 tonnes). Only seven slugs for the 350 mm gun were recovered and destroyed. Iraq claimed never to have received any design, assistance, materials or equipment for the planned rocket -assisted projectiles. At the end of September 1995 UNSCOM obtained information on an indigenous 600 mm indigenous Iraqi supergun design - evidence that the project had not died with Bull. Ch3 68

24 3.9 Concluding Remarks Ch3 69

25 Topics Equations of Motion: Gravity-free, drag-free space flight: Ideal Rocket Equation, Forces acting on vehicle: Vehicle flight in atmosphere and gravity influence. Effect of propulsion system on vehicle performance, Space Flight: Circular and elliptical orbits, Minimum energy transfer maneuvers: Hohmann transfer, Plane and altitude changes, Perturbations, Mission velocity. Flight Maneuvers, Flight Vehicles: Multistage vehicles, Launch vehicles. Ch3 70

26 Specific Impulse linear Vehicle Perf. vs. Prop. System Effects More energetic propellants, higher chamber pressure, larger Mass Ratio logarithmic Reduce inert mass to increase g-t Loss Reduce burning time during vertical ascent by higher thrust, Use wings to add lift. Aerodynamic Drag: Tailor forebody shape, Reduce cross-sectional area, Reduce base drag. Initial Velocity: Launch near equator, Air-launch. Optimize trajectory to increase propulsive efficiency: c u Ch3 71

27 Flight Maneuvers First and upper stage propulsion, Orbit injection or transferring from one orbit to another, Velocity vector adjustments and minor in-flight corrections, Reentry and landing, Rendezvous and docking, Plane change of flight trajectory, Simple rotation, De-orbiting and disposal of used/spent SC, Emergency or alternative mission. Ch3 72

28 3.10 SUPPLEMENT Ch3 73

29 Characteristic Data for Heavenly Bodies Ch3 74

30 Properties of s Standard Atmosphere Ch3 75

31 Earth s Atmosphere Ch3 76