Prof. Claudio Bruno University of Rome Prof. Paul Czysz St. Louis University The Future of Space Depends on Dependable Propulsion Hardware for Non-Expendable.

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

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?

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

A 1985 Estimate for the Beginning of the 21st Century

A 1985 Estimate for the Beginning of the 21st Century

Circa 1985

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

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

VDK-CzyszSizing SystemIdentifies theSolution Spacefor theIdentifiedRequirements

VDK-CzyszSizing SystemIdentifies theSolution Spacefor theIdentifiedRequirements

Where Design ParametersConverge Identifying theSolution Space

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

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

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 ?

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

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

Aerospatiale Mig/Lozinski 50-50

Sänger Daussalt

MAKS Canadian Arrow

Since The 1960’ sThereWereAnd AreManyGoodDesigns

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

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

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

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

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

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

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

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

What’s Wrong with This Picture ???What’s Wrong with This Picture ???

Circa 1985

No Change in the past 40 years !!

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

Can This Be Our Future Infrastructure ? Can This Be Our Future Infrastructure ?

We Need a Nuclear Electric ShuttleWe Need a Nuclear Electric Shuttle

V. Gubonov NPO EnergiaBonn 1972

The Moon Can Be A Development Site for Both Moon & Mars Hardware

The Moon Can Be A Development Site for Both Moon & Mars Hardware

Moon or MarsConditions are similar

This is only a transient visit

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

Cape Verde on Victoria Crater

This is Not Similar

the Moon

Cape Verde on Victoria Crater

This is Not Similar

the Moon

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

We Seem to be Trapped by Chemical PropulsionWill We Lead or Follow ?

We Seem to be Trapped by Chemical PropulsionWill We Lead or Follow ?

Professor Claudio BrunoWill Now TakeUs Beyond MarsToward theHeliopause

Nuclear Propulsion - Present/future interplanetary Nuclear Propulsion - Present/future interplanetary missionsmissions

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

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

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

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 ?

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

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

39

NP Nuclear Energy

mass energy

m mc2

depends on fundamental forces

Nuclear Propulsion - Einstein’s EquationNuclear Propulsion - Einstein’s Equation

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)

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

42

Nuclear Propulsion Isp

Isp/c as function of : the limit Isp = speed of light !

Nuclear Propulsion - IspNuclear Propulsion - Isp

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)

44

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

45

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

46

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

47

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

48

NTR – US Developments (1954-1972)

[M.Turner, “Rocket and Spacecraft Propulsion”, 2005]

Nuclear Propulsion - NTR ApplicationsNuclear Propulsion - NTR Applications

49

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

50

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

51

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

52

52

NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERTravel Time is Still Greater Than One YearTravel Time is Still Greater Than One Year

53

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)

53

NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERDelta V versus PowerDelta V versus Power

54

54

NEP: Applied to ORBIT TRANSFERNEP: Applied to ORBIT TRANSFERPropellant Consumption DominatesPropellant Consumption Dominates

Propellant and Crew Consumables

Propellant

55

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

56

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

57

0

2

4

6

8

10

12

14

16

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

58

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

59

59

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.