INTERSTELLAR PRECURSOR MISSIONS USING ADVANCED DUAL-STAGE ION PROPULSION SYSTEMS

Transcription

1 INTERSTELLAR PRECURSOR MISSIONS USING ADVANCED DUAL-STAGE ION PROPULSION SYSTEMS David G Fearn, 23 Bowenhurst Road, Church Crookham, Fleet, Hants, GU52 6HS, UK dg.fearn@virgin.net Roger Walker Advanced Concepts Team, ESA/ESTEC, Keplerlaan 1, 2201 AZ Noordwijk, Holland Roger.Walker@esa.int ESTEC 21 February

2 Contents Objectives of the study Ion thruster principles of operation Ion thruster scaling Feasible objectives Modified ion extraction and acceleration system High SI and high power operation Performance evaluation Power source Typical mission requirements Implications for spacecraft and ion propulsion system Results of parametric study Description of experiments at ESTEC DS4G ion thruster Results Conclusions 2

3 Objectives Demonstrate that the use of gridded ion thrusters enables very high total impulses to be achieved at ultra-high values of specific impulse Requirements for interstellar precursor missions can be met at moderate cost Useful data return can be achieved within about 25 to 30 years Avoid gravity-assist manoeuvres to keep launch windows flexible Include launch to low Earth orbit to minimise cost Confirm that high thrust density and SI can be achieved Based initially on experimental results from CTR particle injector programmes Produce parametric data to aid future mission design Conduct experimental evaluation of the 4-grid ion extraction and acceleration concept Demonstrate that the two processes can be separated Assess performance capabilities 3

4 Gridded Ion Engine T6 Thruster (QinetiQ) 4

5 Ion Extraction Grid Configuration Triple-grid configuration, indicating the effects of the sheath shape on ion trajectories 5

6 What is Feasible? Authors (Propell) CTR Devices Okumura (H) Grid Size (cm) Open Area Ratio (%) Beam Energy (kev) Beam Current (A) Beam Power (kw) Beam Diverg (deg) Current Density (ma/cm 2 ) Thrust (N) SI (s) 10 dia ,000 Ohara* (H) 12 dia ,000 Menon (H) 18 dia ,000 Martin** (H) ,000 Ion Thrusters Wilbur (Xe) ? ,900 Wilbur (Ar) ? ,000 * Design study ** Production devices from the UKAEA Culham Laboratory Derived from experiments with Ar 6

7 Modified Ion Extraction Grid Configuration Schematic diagram of 4-grid system as used in CTR injection machines. The ion extraction process is separated from the acceleration process. The ion beam current density can be independent of the SI. 7

8 Governing Equations (1) 8

9 Governing Equations (2) Maximum Values of Major Parameters 9

10 Performance Predictions 20 cm beam diameter thruster with 4-grid system Xe propellant 85% propellant utilisation efficiency Flat plasma density profile for long grid life 5 kv ion extraction Net Beam Accelerating Potential (kv) Maximum thrust, 50% perveance (N) Thrust density (mn/cm 2 ) Beam current (A) Input power at 50% Perveance (kw) SI (s) ,100 13,600 15,700 19,300 24,900 29,400 Power density (W/cm 2 )

11 Power Source Must be nuclear fission No other alternative is viable for deep space, high energy missions Solar radiation intensity is negligible Based study on Topaz 2 reactor Selected for cancelled NEPSTP mission 135 kw thermal, 6 kw output from thermionic diodes Mass about 1 tonne, including radiator, shielding, etc Turbo-machinery can raise output to up to 50 kw Radiator mass reduced considerably Assume 22 to 27 kg/kw Consistent with parametric studies of future nuclear technology Power output must be consistent with long life Need to de-rate for long missions 11

12 Reactor Specific Mass 12

13 Spacecraft Schematic Diagram Based on Nuclear Electric Propulsion Space Test Program Mission 13

14 Typical Mission Requirements Mission Typical V (km/s) Minimum SI (s) Nearby planets Outer planets AU ,000 10,000 AU ,000 Interstellar 30, Interstellar precursor mission examples; Kuiper Belt, AU Heliopause, 100 AU Gravitational lens focus of Sun, 550 AU 14

15 Mission Implications for Propulsion System Launch to nuclear-safe altitude of 5000 km Spiral orbit-raise manoeuvre to Earth escape, using NEP V = 5.9 km/s Similar spiral orbit raise manoeuvre to escape from Sun s gravitational field V = 29.8 km/s; total escape V = 35.7 km/s (ignores gravity losses) Continue to accelerate away from the Sun Ultra-high SI required to minimise propellant load and launch mass Assumptions: Tankage mass is 6% of propellant mass Basic spacecraft mass, including payload, is 5 tonnes Power system mass is 25 kg/kw, based on Topaz 2 technology Reactor output selected to match thruster requirements at 0.5 N Thruster life is 20,000 hours; gives number required Mass of thrusters and associated components scaled from existing devices, assuming constant 0.5 N thrust 15

16 Launch Mass as Function of SI Launch Mass (tonnes) M sc = 5 tonnes T = 0.5 N α n = 25 kg/kw M ps = 156 kg Optimum SI = 20,000 s V B = 35 kv P T = 60 kw One 20 cm thruster required Launch mass as low as 8.1 tonnes Specific Impulse (1000 s) 16

17 Input Power and Beam Accel Potential vs SI Input Power (kw) Power (kw) Beam Accel Potential (kv) Beam Accel Potential (kv) Optimum SI = 20,000 s V B = 35 kv P T = 60 kw T = 0.5 N α n = 25 kg/kw Specific Impulse (1000 s) 0 17

18 Nominal Distance Achieved in 25 yr vs SI 150 Distance in 25 yr (AU) M sc = 5 tonnes T = 0.5 N α n = 25 kg/kw M ps = 156 kg 140 AU exceeded Specific Impulse (1000 s) 18

19 Time Required to Achieve Escape V vs SI Time to achieve escape (yr) M sc = 5 tonnes T = 0.5 N α n = 25 kg/kw M ps = 156 kg Time well below specified 25 yr Optimum time does not equate to optimum SI Specific Impulse (1000 s) 19

20 Effect on Optimum SI of Basic Spacecraft Mass 25 Optimum SI (1000 s) T = 0.5 N α n = 25 kg/kw M ps = 156 kg Optimum SI increases with M sc Payload Mass (tonnes) 20

21 Nominal Distance Achieved in 25 yr and Time Required to Achieve Escape V vs M sc Optimum SI Assumed Distance in 25 yr (AU) Distance Time Time to achieve escape (yr) T = 0.5 N α n = 25 kg/kw M ps = 156 kg Performance much improved if M sc can be reduced Distance can exceed 200 AU Payload Mass (tonnes) 0 21

22 Effect of Power Source Specific Mass Launch Mass vs SI Launch Mass (Tonnes) kg/kw 25 kg/kw 35 kg/kw M sc = 5 tonnes T = 0.5 N M ps = 156 kg Optimum SI moves to higher value as α n decreases Optimum is 24,000 s at 15 kg/kw Specific Impulse (1000 s) 22

23 Effect of Nominal Thrust Launch Mass and Input Power vs Thrust Mass Power M sc = 5 tonnes SI = 20,000 s Launch Mass (tonnes) Input Power (kw) M ps variable α n = 25 kg/kw Both parameters increase linearly with thrust 13.5 tonnes at 2 N 240 kw at 2 N Thrust (N) 23

24 Effect of Nominal Thrust Distance Achieved in 25 yr and Time to Reach Escape V vs Thrust Nominal distance in 25 yr (AU) Distance Time Time to achieve escape (yr) M sc = 5 tonnes SI = 20,000 s M ps variable α n = 25 kg/kw 300 AU exceeded at 2 N 7 yr to reach escape Thrust (N) 5 24

25 Experimental Programme at ESTEC Objectives Establish the validity of the physical concepts underlying the 4-grid proposal Demonstrate the separation of the ion extraction and ion acceleration processes Method Utilise simple thruster based on RF discharge for ion production Extract ions from central region of the discharge chamber only to avoid complications caused by plasma density variations over the grid Procedure Start with simple single aperture (1 mm diameter in screen grid) Extend to 43 holes of 1 mm diameter In each case, gradually increase total applied potential to 30 kv Optimise ion extraction potential Measure ion beam characteristics 25

26 Proof-of-Concept Experiment GSP contract established with Plasma Research Laboratory, Australian National University (ANU) in June 2005 ANU designed, developed and delivered a lab model prototype of the Dual-Stage 4-Grid (DS4G) ion thruster to ESA specifications in 4 months DS4G thruster was tested in the ESTEC EPL CORONA vacuum facility during November 2005 with support from ANU and EPL Beam diagnostic tools: Electrostatic wires (beam divergence) Faraday cup (ion current density) Langmuir probe (beam plasma potential, T e, n e ) Tests very successful Extra tests with new grid design planned for May 2006 to further improve performance 26

27 DS4G Thruster Design 5 cm diameter RF discharge chamber ( W) Xenon gas 4 grids with 43 apertures, 2.3 cm beam diameter Grid open area ratio 8% Total beam potential up to 30 kv 27

28 DS4G Thruster in Operation 28

29 Measured Performance Total beam potential up to 30 kv Specific impulse up to 19,200 seconds* Thrust up to 2.85 mn Beam power W Thrust density 8.6 mn/cm 2 (open area)* Power density 740 W/cm 2 (open area)* Propellant utilisation efficiency up to 90% Total efficiency up to 36 % Direct ion impingement on electrodes <1% Beam divergence 2.5 degrees* * World bests with Xe propellant 29

30 Potential Applications Low power, high V, small robotic missions RTG-based 1000kg-class spacecraft to the edge of the solar system e.g. orbiters for Saturn, Neptune, Pluto, KBO rendezvous, local interstellar medium High power, high V, large robotic missions Nuclear reactor-based robotic spacecraft to the outer solar system, with possible sample return capability (200 kw thruster, 1-2 N thrust) Moon/Mars reusable NEP or SEP cargo tugs ( kw thruster) Very high power, high V human missions Nuclear reactor-based multi-mw advanced human Mars spacecraft (small cluster of 1 MW thrusters each of cm diameter) 30

31 Conclusions A 4-grid ion extraction and acceleration system permits very high values of SI to be achieved from a gridded ion thruster Power and thrust densities are of interest for interstellar precursor missions Typical maximum values for a 20 cm beam diameter thruster at 30 kv are 5.8 N thrust, 600 kw input power and an SI of 19,300 s A parametric study has shown that distances from the sun of >300 AU are feasible within 25 yr, with a launch mass of less than 10 tonnes The time to achieve escape velocity can be as little as 7 yr Such a performance is only possible using a nuclear fission power source, with turbomachinery employed for power conversion The preliminary experiments at ESTEC have verified the 4-grid principle Previously unattainable performance data have been achieved These include an SI of 19, 200 s, a thrust density of 8.6 mn/cm 2, and a beam divergence of