The Future of Space Depends on Dependable Propulsion Hardware for Non-Expendable Systems

Prof. Claudio Bruno University of Rome Prof. Paul Czysz St. Louis University

Ad Astrium Possible?

What opportunities have we rejected?

How far can we travel with our hardware capabilities?

What do we need in terms of hardware performance to travel farther within human organizational interest?

Prof. Bruno

Focus on exploring Beyond LEO

Outer Planets Kuiper Belt Heliosphere Prof. Czysz

Focus on LEO, GSO, and Lunar support as Recommended by Augustine Committee

Earth-Moon Inner Planets

A 1985 Estimate for the Beginning of the 21

st

Century

Circa 1985

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 ?

Star Clipper FDL-7MC M=12 Cruise 8 flts/yr For 10 yr 176H Combined Cycle LACE 42 flts between Overhaul P&W XLR-129 SERJ

VDK-Czysz Sizing System Identifies the Solution Space for the Identified Requirements Where Design Parameters Converge Identifying the Solution Space



Necessary Volume and Size for SSTO Blended Body Convergence

Blended Body

ICI

 ICI 

ppl

  Propulsion Index/Structural Index

MR W str



ppl S wet

Propellant density

MR

 Mass ratio

W str S wet

  Structural weight Vehicle surface area Impractical Solution area

Delineates the possible from the not possible

Little Difference in Empty Weight, A Significant Difference in Gross Weight

Practical Solution Space within Industrial Capability about 1/5 the Total Possible

The Solution Space for Four Configuration Concepts Identifies Configuration Limitations

ft 2

Why was Delta Clipper A Circular Cone ?

Even an All Rocket TSTO Has More Versatility, Flexibility & Payload Volume Than a SSTO A TSTO is One-Half the Mass

Individual components 1 st Stage

Staging Above Mach 10 Minimizes TSTO System Weight

TSTO system Dwight Taylor McDonnell Douglas Circa 1983 Toss-Back is all metal toss-back booster staging at Mach 7 is low cost, fully recoverable and sustained use at acceptable mass

Since The 1960’ s There Were And Are Many Good Designs Aerospatiale Sänger MAKS Mig/Lozinski 50-50 Daussalt Canadian Arrow

Cargo ISS Crew

As a First Step We Can Have a Versatile, Flexible, Recoverable and Reusable Rocket System

From McDonnell Douglas Astronautics, Huntington Beach, circa 1983 It can be a rocket and does not have to be an ejector rocket/scramjet

40% penalty Unless the WR is Less Than 5.5 HTO is an Unacceptable Penalty HTO is not a Management Option !!

Airbreathing Option Pays At Speeds Less Than 14,500 ft/sec

Confirmed by A Blue Ribbon Panel Headed by Dr. B. Göthert in Circa 1964 After Reviewing Available Data

OWE Solution Spaces Overlap. Marginal Difference in OEW

LACE Offers An Existing Rocket Benefit Almost Equal to a Combined Cycle

Popular Choice not the Better Choice

1 st Stage Propulsion Gross Weight (ton) 1 st Stage Stage Weight (ton) Propellant Wt. (ton) Engine Weight (ton) Dry Weight (ton) 2 nd Stage Stage Weight (ton) Propellant Wt. (ton) Engine Weight (ton) Dry Weight (ton) Turbo-Ramjet 393 283 83.2

60.5

200 109 81.6

7.0

20.3

Ejector-Ramjet 261 142 45.5

7.3

96.1

118 87.9

7.0

23.5

Thrust @ Mach 6.7 compared ≈ 1 ≈ 0.25

to thrust @ takeoff

Expendable Sustained Use

10 year Operational Life, 30,000 lb payload, Up to 10 Flights/year per Aircraft for Four Propulsion Systems

By H. D. Froning And Skye Lawrence Circa 1983 Sustained Use

LLC Constant

Cost Data is Consistent, Fly More Often With Sustained Use Aircraft

$10 5 $/lb = 46951. * FR – 0.638

By H. D. Froning And Skye Lawrence Circa 1983 $10 4 $/lb = 77094. • FR – 0.985

$10 3 $10 2 1.0

B-747 flying at same rate and payload as shuttle Rocket Min. AB M ax. AB Current exp.

10 FR Flight Rate Flights/Year 10 2

It’s the FLIGHT RATE, not technology

100,000 10,000 Shuttle O’Keefe Rocket Airbreather < M 10 Airbreather > M 10 Scaled 747 Operations Current Partial Reusable 1,000 100 10 5 B747’s Operated At Same Schedule And payload As The Space Shuttle Charles Lindley, Jay Penn Aviation Week and Space Technology , June 15, 1998 The Aerospace Corp. Database 1 1.0

10 100 Flights/year 1,000 10,000 100,000 1,000,000 10,000,000

What’s Wrong with This Picture ???

No Change in the past 40 years !!

Circa 1985

Augustine 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

AMERICA

October 2009 Page 19

Can This Be Our Future Infrastructure ?

We Need a Nuclear Electric Shuttle

V. Gubonov NPO Energia Bonn 1972

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

Moon or Mars Conditions are similar This is only a transient visit

Moon-Mars Human Infrastructure Needs 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

Chemical Propulsion is a Poor Option to Mars 250 200 150 100 50 0

Hypergolic H2/O2

Propulsion Systems

Nuclear Rubbia

Mars

We Seem to be Trapped by Chemical Propulsion Will We Lead or Follow ?

Nuclear Propulsion - Present/future interplanetary missions

Professor Claudio Bruno Will Now Take Us Beyond Mars Toward the Heliopause

Nuclear Propulsion - Times and distances of present/future interplanetary missions Manned

: constrained by physical/psychological support air, victuals cosmic & solar radiation, flares bone/muscle mass loss enzymatic changes, …?

Unmanned

: public support, apathy @ > 1-2 years: funding difficult To reduce constraints, risks, and ensure public (financial) support

faster missions with less mass (cost ~ mass)

33

Nuclear Propulsion - Times and distances with Acceleration 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 distance 1/2 dist time time V 1/2 V 1/2 /c WR 1/2 1/100 4.05E+09 2.02E+09 0.258 94.31 799.13 0.43% 7.52E+77 1/10,000 4.05E+09 2.02E+09 2.582 943.14 79.91 0.043% 1.25E+07 Boost-coast 4.05E+09 2.02E+09 11.284 4,121 18.29 0.010% 10.28 “g” miles miles years days km/sec % light speed

34

Nuclear Propulsion - Times and Isp

At a = 10 -2 g, trip is fast, but: mass ratio is significant. What compromises between mass ratio and time ? Nuclear propulsion looks feasible if Isp can be raised:

Jupiter Saturn Uranus Neptune Kuiper Belt Pluto Kuiper Belt Heliopause Isp (sec) WR years 2.69 4.92 8.14 11.15 11.13 13.75 16.29 27.86 459 10.70 years 1.70 3.12 5.16 7.07 7.06 8.72 10.34 17.67 1,100 7.23 years 0.793 1.45 2.40 3.29 3.29 4.06 4.81 8.22 4,590 3.38

Increasing Isp Reduces Transit Time and Weight Ratio 35

If J = specific energy (energy/unit mass) 1-D, ideal, propellants acceleration:

J = (1/2) V e 2

thus:

Isp = V e = (2J) 1/2

Ve = exhaust velocity = Isp [m/s]  to increase Isp, J must be increased

much more

36

F aster missions, lower mass consumption feasible with / if  non-zero acceleration  not boost-coast  higher Isp

Isp = V e = (2J) 1/2



thrust power ~ Isp 3 = (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

37

Nuclear Propulsion - Energy Density in NP - Energy Density in Chemical Propulsion Chemical propellants

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 ~ 2

J



Must increase J by orders of magnitude



Nuclear energy

38



mass energy

m

a

mc 2

 a

depends on fundamental forces

39

Nuclear Propulsion Potential Energy

Compare alphas and energies:

 a and energy density J ( J = [E/m] = a c 2 )   No known a Even a between 3.75 x 10 -3 and 1 = 1 produces not directly useable energy (e.g., g rays) 40

Calculate Isp:    Assume ideal expansion (to pe=0): Isp = V Obtaining Ve is a 3-stage process: Pot. Energy Microenergy of matter Thermalization Orderly bulk motion V ( ( e.g., Vibr., Transl., Ionization, n, e , α + ) e ≡ Possible addition of inert mass, Mp from relativistic energy balance: (equilibrium) 2 a ) 2  V 1 2 (for short)

m o

(1  a 1 

V c

2 at V = Ve )

V

2 2  1 2

Mp

1 

o V V c

2 2 2 Plot normalized specific impulse, Isp/c = V /c = Ve /c: 41

Isp/c as function of

a

: the limit Isp = speed of light !

42

 Satisfies both    F·Isp = P , thrust power = η tot x P reactor  ( m = total mass rate ejected )  1 2 grows slowly with P R, ~ reactor cost  Thus, in terms of inert mass addition, or μ  Where z: F = α 

m

0 η tot     

=

1 : unreacted fuel also ejected  = 0 : unreacted fuel stays inside reactor  Thrust may be written  0  

Limit thrust

 2  1 2  η Φ z, α, μ, tot V c

Amplification factor

43

Nuclear Propulsion Thrust Power P P Let’s look at the power needed by F:   e e  3 1-f 2    P = F · Isp = F · V 

Trade off between F and I sp

 P scales with V 3 : ‘high’ thrust (‘fast’) missions need ‘much larger’ P, affordable ONLY with nuclear power 44

Nuclear Propulsion - How to Utilize Nuclear Power

Two strategies: NTR (Nuclear Thermal Rockets): expand hot fluid, as in chemical rockets. E.g., with H 2 and max T = 3000K  Isp ~ 1000 s, thermal efficiency ≈ 1 (all heat absorbed by H 2 ). Bulk power density ~ 10 -3 to 10 -1 kg/kW. NTR may be very compact, e.g., with 242 Am 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 ~ 10 5 – 10 6 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 45

Nuclear Propulsion - Application Strategies

Schematics of NTR – Nuclear Thermal Rocket

Figure 7-6

:

Conceptual scheme of a Nuclear Thermal Rocket (Bond, 2002)

46

Nuclear Propulsion - Application Strategies

Schematics of NER – Nuclear Electric Rocket

Figure 7-7

:

Conceptual scheme of a Nuclear-Electric Rocket. Note the mandatory radiator (Bond, 2002)

47

Nuclear Propulsion - NTR Applications

NTR – US Developments (1954-1972)

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

48

Nuclear Propulsion - NTR Applications

NTR – US Developments (1954-1972)

The Phoebus IIA solid-core nuclear reactor on its Los Alamos test stand (Dewar, 2004 )

49

Nuclear Propulsion - Application Strategies

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) 50

Nuclear Propulsion - Comparisons

 Must set ground rules (otherwise, apples & pears)  Here: based on I tot,s = (I sp t operation )/(M P + m) ~ Isp 3 η tot /P R I tot,s is a distance traveled/unit ‘fuel’ mass, as in cars  Normalize I tot,s using I tot,s of LOX/LH 2 : this ratio is the ‘ performance Index, I’: 51

NEP: Applied to ORBIT TRANSFER Travel Time is Still Greater Than One Year

52 52

NEP: Applied to ORBIT TRANSFER Delta V versus Power

ΔV (km/s) NEP Power (MWe) 100 150 200 300 Total ΔV (km/s) 86.2

103.2

106.7

114.8

POWER (Mwe) MASS: 120 to160 ton Compared with CP total ΔV is 406.76% to 574.9% higher

53 53

NEP: Applied to ORBIT TRANSFER Propellant Consumption Dominates

Propellant and Crew Consumables Propellant

54 54

Power to Travel 73 AU Distance

1.E+07 1.E+06 1.E+05 1.E+04 1.E+03 1.E+02

Kuiper Belt

10 [Mo, kg] 100 1000 10000 100000 1.E+01 50 100 150 200 250 300 350

Isp [km/s] - 73 AU / 8 years

400 450 500

Power as function of Isp; 8-year mission and initial mass M 0 as parameter order of magnitude more power than 20 year mission

55

Power to Travel 73 AU Distance

1.E+07 1.E+06 1.E+05 1.E+04 1.E+03 1.E+02 1.E+01 50 100 150 200 250 300 350

Isp [km/s] - 73 AU / 20 years

400 450 500

Power as function of Isp; 20-year mission and initial mass M 0 as parameter Kuiper Belt

10 [Mo, kg] 100 1000 10000 100000 56

16 14 12 10 8 6 4 2 0 0 1 10 100 a

[kW/kg] - ML/M0 = 0.1

1000 10000 45 40 35 30 25 20 15 10 5 0 0

100 AU

50 [km/s] 150 250 350 450

Power to Travel to the Heliopause 100 AU Distance for Two Travel Times 100 AU

50 [km/s] 150 250 350 450 1 10 100 a

[kW/kg] - ML/M0 = 0.6

1000 10000 57

80 70 60 50 40 30 20 10 0 0

540 AU

50 [km/s] 150 250 350 450

540 AU Distance to the Sun Focal Point for Two Travel Times

1 10 100 a

[kW/kg] - ML/M0 = 0.1

1000 10000 250 200 150 100 50 0 0

540 AU

50 [km/s] 150 250 350 450 1 10 100 a

[kW/kg] - ML/M0 = 0.6

1000 10000 58

Nuclear 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.

 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 59 59