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![Page 1: Prof. Claudio Bruno University of Rome Prof. Paul Czysz St. Louis University The Future of Space Depends on Dependable Propulsion Hardware for Non-Expendable.](https://reader031.fdocuments.us/reader031/viewer/2022012918/5514f9e7550346b0338b634d/html5/thumbnails/1.jpg)
Prof. Claudio BrunoUniversity of Rome
Prof. Paul CzyszSt. Louis University
The Future ofSpace Depends on Dependable
PropulsionHardware for
Non-Expendable Systems
The Future ofSpace Depends on Dependable
PropulsionHardware for
Non-Expendable Systems
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Ad AstriumPossible?
Ad AstriumPossible?
What opportunitieshave we rejected?
How far can we travelwith our hardware capabilities?
What do we need in terms of hardwareperformance to travelfarther within humanorganizational interest?
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Earth-MoonInner Planets
Outer PlanetsKuiper BeltHeliosphere
Prof. Bruno
Prof. Czysz
Focus on LEO, GSO, and Lunar support as Recommended by Augustine Committee
Focus on exploringBeyond LEO
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A 1985 Estimate for the Beginning of the 21st Century
A 1985 Estimate for the Beginning of the 21st Century
Circa 1985
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Space and Atmospheric Vehicle Development Converge, So the Technology of High Performance Launchers Applies to
Airbreathing Aircraft, Aeronautics and Astronautics 1971
Space and Atmospheric Vehicle Development Converge, So the Technology of High Performance Launchers Applies to
Airbreathing Aircraft, Aeronautics and Astronautics 1971
Buck, Neumann & Draper were Correct in 1965
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What If These 1960’s Opportunities Were Not Missed ?
What If These 1960’s Opportunities Were Not Missed ?
Star Clipper M=12 Cruise FDL-7MC
176H SERJCombined Cycle
LACE8 flts/yrFor 10 yr 42 flts between
Overhaul P&W XLR-129
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VDK-CzyszSizing SystemIdentifies theSolution Spacefor theIdentifiedRequirements
VDK-CzyszSizing SystemIdentifies theSolution Spacefor theIdentifiedRequirements
Where Design ParametersConverge Identifying theSolution Space
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Necessary Volume and Size for SSTO Blended Body Convergence
Necessary Volume and Size for SSTO Blended Body Convergence
ImpracticalSolution area
BlendedBody
ICI Propulsion Index/Structural Index
ICI MR ppl
Wstr Swet
ppl Propellant density
MR Mass ratio
Wstr Structural weight
Swet Vehicle surface area
Delineates the possible from the not possible
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Little Difference in Empty Weight,A Significant Difference in Gross Weight
Little Difference in Empty Weight,A Significant Difference in Gross Weight
0.20
0 40,000 80,000 120,000
Operational Empty Weight OEW (lbs)
1,200,000
1,000,000
800.000
600.000
400,000
200,000
0
SSTO Solution SpaceRocket
M=12 Combined Cycle
0.20
0.16
0.12
0.80
0.10
0.063
0.16 0.12 0.800.063
1
2.5
0
5
7
10
Payload(tons)
tau
Payload(tons)
tau
12.5
0
57
10
20 30 40 50 60(tons) 10
500
100
(tons)
GW
G
ross
We
igh
t (
lbs)
Practical Solution Space within Industrial Capability about 1/5 the Total Possible
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The Solution Space for Four Configuration Concepts Identifies Configuration Limitations
The Solution Space for Four Configuration Concepts Identifies Configuration Limitations
ft2
Why was Delta ClipperA Circular Cone ?
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Even an All Rocket TSTO
Has MoreVersatility,Flexibility& Payload
Volume Than a SSTO
A TSTO is One-Half the Mass
Even an All Rocket TSTO
Has MoreVersatility,Flexibility& Payload
Volume Than a SSTO
A TSTO is One-Half the Mass
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Staging Above Mach 10
Minimizes TSTO System Weight
Staging Above Mach 10
Minimizes TSTO System Weight
Toss-Back is all metaltoss-back boosterstaging at Mach 7is low cost, fullyrecoverable andsustained useat acceptable mass
TSTO systemDwight TaylorMcDonnell DouglasCirca 1983
Individual components 1st Stage
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Aerospatiale Mig/Lozinski 50-50
Sänger Daussalt
MAKS Canadian Arrow
Since The 1960’ sThereWereAnd AreManyGoodDesigns
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As a First Step We Can
Have aVersatile,Flexible,
Recoverableand
Reusable RocketSystem
As a First Step We Can
Have aVersatile,Flexible,
Recoverableand
Reusable RocketSystem
From McDonnell Douglas Astronautics, Huntington Beach, circa 1983
It can be a rocket and does not have to be an ejector rocket/scramjet
Cargo ISS Crew
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Unless the WR is Less Than 5.5 HTO is an
Unacceptable Penalty
Unless the WR is Less Than 5.5 HTO is an
Unacceptable Penalty
HTO is not aManagement
Option !!
40% penalty
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AirbreathingOption PaysAt SpeedsLess Than
14,500 ft/sec
AirbreathingOption PaysAt SpeedsLess Than
14,500 ft/sec
Confirmed byA Blue RibbonPanel Headed byDr. B. Göthert inCirca 1964After ReviewingAvailable Data
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LACE Offers AnExisting RocketBenefit Almost
Equal to a Combined Cycle
LACE Offers AnExisting RocketBenefit Almost
Equal to a Combined Cycle
OWE Solution SpacesOverlap. MarginalDifference in OEW
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Popular Choice not the Better ChoicePopular Choice not the Better Choice
1st Stage Propulsion
Turbo-Ramjet Ejector-Ramjet
Gross Weight (ton)
393 261
1st Stage
Stage Weight (ton)
283 142
Propellant Wt. (ton)
83.2 45.5
Engine Weight (ton)
60.5 7.3
Dry Weight (ton) 200 96.1
2nd Stage
Stage Weight (ton)
109 118
Propellant Wt. (ton)
81.6 87.9
Engine Weight (ton)
7.0 7.0
Dry Weight (ton) 20.3 23.5
Thrust @ Mach 6.7 compared ≈ 1 ≈ 0.25 to thrust @ takeoff
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10 year Operational Life, 30,000 lb payload, Up to 10 Flights/year per Aircraft for FourPropulsion Systems
10 year Operational Life, 30,000 lb payload, Up to 10 Flights/year per Aircraft for FourPropulsion Systems
By H. D. FroningAndSkye LawrenceCirca 1983
By H. D. FroningAndSkye LawrenceCirca 1983
Expendable
Sustained Use
Sustained Use
LLC Constant
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Cost Data is Consistent, Fly More OftenWith Sustained Use Aircraft
Cost Data is Consistent, Fly More OftenWith Sustained Use Aircraft
$/lb = 46951. * FR– 0.638
Current exp.
Max. AB
Min. AB
Rocket
B-747 flying at samerate and payload as shuttle
$/lb = 77094. • FR– 0.985
FR Flight Rate Flights/Year
10 102$102
$103
$104
$105
1.0
Cos
t of
Pay
load
to O
rbit
( $
/lb )
•
By H. D. FroningAndSkye LawrenceCirca 1983
By H. D. FroningAndSkye LawrenceCirca 1983
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Current Partial Reusable
Scaled 747 Operations
Airbreather > M 10
Airbreather < M 10
Rocket
1
10
100
1,000
10,000
100,000
Cos
t of
Pay
load
(
$/lb
)
Flights/year10,000,0001,000,000100,00010,0001,000100101.0
Aviation Week and Space Technology,June 15, 1998The Aerospace Corp. Database
It’s the FLIGHT RATE, not technologyIt’s the FLIGHT RATE, not technology
Charles Lindley,Jay Penn
5 B747’s OperatedAt Same ScheduleAnd payload AsThe Space Shuttle
ShuttleO’Keefe
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What’s Wrong with This Picture ???What’s Wrong with This Picture ???
Circa 1985
No Change in the past 40 years !!
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Augustine CommitteeAugustine Committee
Review of Human Spaceflight Plans Committee expressed an eagerness with a concept that with Werner von Braun originated in the 1950’s – orbital refueling.
AEROSPACE AMERICAOctober 2009Page 19
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Can This Be Our Future Infrastructure ? Can This Be Our Future Infrastructure ?
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We Need a Nuclear Electric ShuttleWe Need a Nuclear Electric Shuttle
V. Gubonov NPO EnergiaBonn 1972
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The Moon Can Be A Development Site for Both Moon & Mars Hardware
The Moon Can Be A Development Site for Both Moon & Mars Hardware
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Moon or MarsConditions are similar
This is only a transient visit
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Moon-Mars Human InfrastructureNeeds to be Proven by Sustained
Applications, First on the Moon Then Mars
Moon-Mars Human InfrastructureNeeds to be Proven by Sustained
Applications, First on the Moon Then Mars
We need to lift Habitats, Food, Water, Green Houses and Soil Handling Equipment In Addition to People to confirm long term hardware viability
RTV powered Automatic Greenhouse With 10 year operational life
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Cape Verde on Victoria Crater
This is Not Similar
the Moon
Cape Verde on Victoria Crater
This is Not Similar
the Moon
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Chemical Propulsion is a Poor Option to MarsChemical Propulsion is a Poor Option to Mars
HypergolicH2/O2
NuclearRubbia
Mars
0
50
100
150
200
250
Tra
ns
it t
ime
(d
ay
s)
Propulsion Systems
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We Seem to be Trapped by Chemical PropulsionWill We Lead or Follow ?
We Seem to be Trapped by Chemical PropulsionWill We Lead or Follow ?
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Professor Claudio BrunoWill Now TakeUs Beyond MarsToward theHeliopause
Nuclear Propulsion - Present/future interplanetary Nuclear Propulsion - Present/future interplanetary missionsmissions
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33
Nuclear Propulsion - Times and distances of Nuclear Propulsion - Times and distances of present/future interplanetary missionspresent/future interplanetary missions
Manned: constrained by physical/psychological support
To reduce constraints, risks, and ensure public (financial) support
faster missions with less mass (cost ~ mass)
air, victuals
cosmic & solar radiation, flares
bone/muscle mass loss
enzymatic changes, …?
Unmanned: public support, apathy @ > 1-2 years: funding difficult
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34
NP - Times and distances with acceleration
Accelerated travel makes tremendous difference in time to destination
However: mass consumption may be forbiddingly high
e.g.: mission to Neptune, chemical propulsion, Isp = 459 s:
acceleration 1/100 1/10,000 Boost-coast “g”
distance 4.05E+09 4.05E+09 4.05E+09 miles
1/2 dist 2.02E+09 2.02E+09 2.02E+09 miles
time 0.258 2.582 11.284 years
time 94.31 943.14 4,121 days
V1/2 799.13 79.91 18.29 km/sec
V1/2 /c 0.43% 0.043% 0.010% % light speed
WR1/2 7.52E+77 1.25E+07 10.28
Nuclear Propulsion - Times and distances Nuclear Propulsion - Times and distances with Accelerationwith Acceleration
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35
At a = 10-2g, trip is fast, but: mass ratio is significant.
What compromises between mass ratio and time ?
Nuclear propulsion looks feasible if Isp can be raised:
years years years Jupiter 2.69 1.70 0.793 Saturn 4.92 3.12 1.45 Uranus 8.14 5.16 2.40 Neptune 11.15 7.07 3.29
Kuiper Belt 11.13 7.06 3.29 Pluto 13.75 8.72 4.06
Kuiper Belt 16.29 10.34 4.81 Heliopause 27.86 17.67 8.22
Isp (sec) 459 1,100 4,590
WR 10.70 7.23 3.38
Increasing Isp Reduces Transit Time and Weight Ratio
Nuclear Propulsion - Times and Isp Nuclear Propulsion - Times and Isp
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36
NP - What it really means ‘to increase Isp’
If J = specific energy (energy/unit mass) 1-D, ideal, propellants acceleration:
J = (1/2) Ve2 Ve = exhaust velocity = Isp [m/s]
thus:
Isp = Ve = (2J)1/2
to increase Isp, J must be increased much more
Nuclear Propulsion - What Increases Isp ?Nuclear Propulsion - What Increases Isp ?
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37
NP - Mission Time and Power
Faster missions, lower mass consumption feasible with / if
non-zero acceleration not boost-coast
higher Isp Isp = Ve = (2J)1/2
thrust power ~ Isp3 = (2J)3/2
faster missions + high Isp = large power
Large mass consumption: driven by low J of chemical propellants
J of Chemical Propellants 4.0 to 10.0 MJ/kg too low
need to find higher energy density materials
Nuclear Propulsion - Mission Time & Nuclear Propulsion - Mission Time & PowerPower
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38
NP - Energy Density in Chemical Propulsion
Max performance improvement with chemical propulsion:with metallic Hydrogen, theoretical Isp ~ 1000-1700 s existence, stability, control of energy release unsolved issues
J increases by O(10) at most, but Isp ~
Must increase J by orders of magnitude Nuclear energy
2J
Nuclear Propulsion - Energy Density inNuclear Propulsion - Energy Density in Chemical propellantsChemical propellants
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39
NP Nuclear Energy
mass energy
m mc2
depends on fundamental forces
Nuclear Propulsion - Einstein’s EquationNuclear Propulsion - Einstein’s Equation
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40
Nuclear Propulsion Potential Energy Compare alphas and energies:
and energy density J ( J = [E/m] = c2 )
No known between 3.75 x 10-3 and 1 Even = 1 produces not directly useable energy (e.g., rays)
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41
Propulsion Isp
Plot normalized specific impulse, Isp/c = V/c = Ve/c:
Assume ideal expansion (to pe=0): Isp = Ve ≡ V (for short)
Obtaining Ve is a 3-stage process:
Calculate Isp:
Pot. Energy Microenergy of matter Thermalization Orderly bulk motion ((e.g., Vibr., Transl., Ionization, n, e-, α+) (equilibrium) at V = Ve
Possible addition of inert mass, Mp
V from relativistic energy balance: 2 2
2 2
2 2
2 2
(1 )1 1(1 )
2 21 1
o oo o
V V
V
m Mpm c m c
c cV
Nuclear Propulsion - IspNuclear Propulsion - Isp
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42
Nuclear Propulsion Isp
Isp/c as function of : the limit Isp = speed of light !
Nuclear Propulsion - IspNuclear Propulsion - Isp
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43
Propulsion Thrust F
Satisfies both F·Isp = P , thrust power = ηtot x Preactor F=Isp m ( m = total mass rate ejected )
12F = P m grows slowly with PR, ~ reactor cost
Thus, in terms of inert mass addition, or μ
1 2
2 2
0 totF = α c η z 1-α 1- V c +μ 1- V cm
Where z: = 1 : unreacted fuel also ejected = 0 : unreacted fuel stays inside reactor
generally ; if only fission/fusion fragments are ejected, μ = 0 F μ
Thrust may be written 0 tot
VF = α m c η Φ z, α, μ,
c
Limit thrust Amplification factor
Nuclear Propulsion - Thrust (F)Nuclear Propulsion - Thrust (F)
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Nuclear Propulsion Nuclear Propulsion
Thrust Power PThrust Power PLet’s look at the power needed by F:
P = F · Isp = F · V
P scales with V3: ‘high’ thrust (‘fast’) missions need ‘much larger’ P, affordable ONLY with nuclear power
Trade off between F and Isp
3 3 2
e eP ρ A c f 1-f
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Two strategies: NTR (Nuclear Thermal Rockets): expand hot fluid, as in chemical rockets. E.g., with H2 and max T = 3000K Isp ~ 1000 s, thermal efficiency ≈ 1 (all heat absorbed by H2). Bulk power density ~ 10-3 to 10-1 kg/kW. NTR may be very compact, e.g., with 242Am fuel, 40 MW from a 300-kg reactor are feasible. NER/NEP (Nuclear Electric Rocket/Propulsion): run hot fluid in a cycle to generate electric power and feed it to an electric thruster (ET), f.i., ion, arcjet, MPD,…
Isp is that of ET: may be ~ 105 – 106 s and higher. Thermal efficiency: 30-50%; ET efficiency: 70-80%; needs space radiator(s). Bulk power density: low, ~ 1/100 of that of NTR
Nuclear Propulsion - How to Utilize Nuclear PowerNuclear Propulsion - How to Utilize Nuclear Power
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Schematics of NTR – Nuclear Thermal Rocket
Figure 7-6: Conceptual scheme of a Nuclear Thermal Rocket (Bond, 2002)
Nuclear Propulsion - Application StrategiesNuclear Propulsion - Application Strategies
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Schematics of NER – Nuclear Electric Rocket
Figure 7-7: Conceptual scheme of a Nuclear-Electric Rocket. Note the mandatory radiator (Bond, 2002)
Nuclear Propulsion - Application StrategiesNuclear Propulsion - Application Strategies
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NTR – US Developments (1954-1972)
[M.Turner, “Rocket and Spacecraft Propulsion”, 2005]
Nuclear Propulsion - NTR ApplicationsNuclear Propulsion - NTR Applications
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NTR – US Developments (1954-1972)
The Phoebus IIA solid-core nuclear reactor on its Los Alamos test stand (Dewar, 2004 )
Nuclear Propulsion - NTR ApplicationsNuclear Propulsion - NTR Applications
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Nuclear propulsion strategies
Nuclear Electric Propulsion
Two main NEP classes: charged species accelerated by:
Coulomb Force (only electric field imposed)
Lorentz’ forces (electric and magnetic field)
Nuclear Propulsion - Application StrategiesNuclear Propulsion - Application Strategies
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Must set ground rules (otherwise, apples & pears)
Here: based on Itot,s = (Isp toperation)/(MP + m) ~ Isp3 ηtot/PR
Itot,s is a distance traveled/unit ‘fuel’ mass, as in cars
Normalize Itot,s using Itot,s of LOX/LH2 : this ratio is the ‘performance Index, I’:
Nuclear Propulsion - ComparisonsNuclear Propulsion - Comparisons
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NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERTravel Time is Still Greater Than One YearTravel Time is Still Greater Than One Year
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NEPNEP
Power (MWe)
Total ΔV
(km/s)
100 86.2
150 103.2
200 106.7
300 114.8
Compared with CP total ΔV is 406.76% to 574.9% higher
MASS: 120 to160 ton
POWER (Mwe)
ΔV (km/s)
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NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERDelta V versus PowerDelta V versus Power
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NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERPropellant Consumption DominatesPropellant Consumption Dominates
Propellant and Crew Consumables
Propellant
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1.E+01
1.E+02
1.E+03
1.E+04
1.E+05
1.E+06
1.E+07
50 100 150 200 250 300 350 400 450 500
Isp [km/s] - 73 AU / 8 years
Po
wer
[W
]
10 [Mo, kg]
100
1000
10000
100000
Power as function of Isp; 8-year mission and initial mass M0 as parameter order of magnitude more power than 20 year mission
Power to Travel 73 AU DistancePower to Travel 73 AU DistancePower to Travel 73 AU DistancePower to Travel 73 AU Distance
Kuiper Belt
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Power to Travel 73 AU DistancePower to Travel 73 AU DistancePower to Travel 73 AU DistancePower to Travel 73 AU Distance
1.E+01
1.E+02
1.E+03
1.E+04
1.E+05
1.E+06
1.E+07
50 100 150 200 250 300 350 400 450 500
Isp [km/s] - 73 AU / 20 years
Po
wer
[W
]
10 [Mo, kg]
100
1000
10000
100000
Power as function of Isp; 20-year missionand initial mass M0 as parameter
Kuiper Belt
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0
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10
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0 1 10 100 1000 10000
[kW/kg] - ML/M0 = 0.1
T [
year
s]
50 [km/s]
150
250
350
450
100 AU
0
5
10
15
20
25
30
35
40
45
0 1 10 100 1000 10000
[kW/kg] - ML/M0 = 0.6
T [
year
s]
50 [km/s]
150
250
350
450
100 AU
Power to Travel to the Power to Travel to the Heliopause Heliopause
100 AU Distance for100 AU Distance for Two Travel TimesTwo Travel Times
Power to Travel to the Power to Travel to the Heliopause Heliopause
100 AU Distance for100 AU Distance for Two Travel TimesTwo Travel Times
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0
10
20
30
40
50
60
70
80
0 1 10 100 1000 10000
[kW/kg] - ML/M0 = 0.1
T [
year
s]
50 [km/s]
150
250
350
450
0
50
100
150
200
250
0 1 10 100 1000 10000
[kW/kg] - ML/M0 = 0.6
T [
year
s]50 [km/s]
150
250
350
450
540 AU
540 AU
540 AU Distance to 540 AU Distance to the Sun Focal Point the Sun Focal Point
forfor Two Travel TimesTwo Travel Times
540 AU Distance to 540 AU Distance to the Sun Focal Point the Sun Focal Point
forfor Two Travel TimesTwo Travel Times
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To enable a future NEP M3, investing in this propulsion technology is necessary. That is an unlikely prospective in the current financial climate, but would spare much time and effort to our future generations.
NTR systems may be the only propulsion enabling quick reaction missions, e.g., to counter unexpected asteroid threats
Nuclear Propulsion ~ Some ConclusionsNuclear Propulsion ~ Some Conclusions
The combination of Isp and power of the Gridded Ion System for a M3 result in predictions for both mass and mission times that are significantly better than with other CP and NTR propulsion systems.
A NEP-powered M3 appears not only feasible, but also more convenient than CP- and likely also NTR-powered missions in terns of cost, besides being the only way to drastically reduce HUMEX travel time and thus GCR dose for the crew.
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