40 kg to LEO: A Low Cost Launcher for

Australia By Nicholas Jamieson

Thesis topic: ◦ Design of a 40kg to LEO launch vehicle with a hypersonic second stage

Supervisors: ◦ Dr Graham Doig (University of New South Wales)

◦ Dr Steven Tsitas (Australian Centre for Space Engineering Research)

Currently no existence of a dedicated launch vehicle designed to transport small payloads of approximately 40kg to a Lower Earth Orbit

Commercial leverage of the 6U CubeSat as a market driver

Through the development of a new low cost launch vehicle scientists and engineers will be afforded a “low cost and timely way to carry out research in space allowing new ideas and technologies to be explored” through a relatively inexpensive means

Source: Tsitas, S. R., & Kingston, J. (2012). 6U CubeSat commercial applications. The Aeronautical Journal, page 189-198.

Major Problem faced in the launch operations of current small payloads:

◦ The launching of CubeSat’s and other small payloads directly depends

upon the availability of much larger launch vehicles

◦ The auxiliary payloads are requested to wait for the main payload to be scheduled for launch which can take anywhere from a few weeks to a few years

Source: Heyman, J. (2009, October). “FOCUS: CubeSats - A Costing + Pricing Challenge” SatMagazine.

“Current technology is being operated close to its theoretical limit with only marginal efficiency improvement achievable”

Image Source: Wikipedia Source: Jazra, T., & Smart, M. K. (2011). Design Methodology for the Airbreathing Second Stage of a Rocket-Scramjet-Rocket Launch Vehicle. AIAA International Space Planes and Hypersonic Systems and Technologies Conference (pp. 1-26). San Francisco: American Institute of Aeronautics and Astronautics.

The major benefit of employing a scramjet as the hypersonic second stage is the significantly higher specific impulse that it provides for the duration of its operation

Source: Daines, R., & Segal, C. (1998). Combined Rocket and Airbreathing Propulsion Systems for Space Launch Applications. Journal of Propulsion and Power, pages 605-612.

No need to carry oxidiser

A rocket-scramjet-rocket propulsion system injects 3 times the inert mass than a purely rocked based system

Flight operability advantages ◦ Increased mission flexibility with regard to the launch window

◦ Improved offset and rendezvous

◦ Lower cost than a pure rocket based propulsion system

Source: Bilardo, V. J., Curran, F. M., Hunt, J. L., Lovell, N. T., Maggio, G., Wilhite, A. W., et al. (2003). The Benefits of Hypersonic Airbreathing Launch Systems for Access to Space. 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference (pp. 1-14). Huntsville, Alabama: American Institute of Aeronautics and Astronautics.

Capability

Operability

Reliability

Economical

Commercially Orientated Design

Source: Nichols, E. E. (1997). The Space Launch Payload Process. The AIAA Journal, 217-224.

Initial design constraints

Assumptions made along the way

Mathematical Methodology

Resultant Sizing Parameters

Three specified initial design constraints:

◦ The launch vehicle must be able to attain an altitude in Lower Earth Orbit of 600 km

◦ The launch vehicle must be capable of transporting a 40kg payload to said altitude

◦ The launch vehicle must be able to be launched into a polar orbit

Source: Google Maps

1. As this is a preliminary sizing, atmospheric drag was neglected in the calculations.

2. It was also assumed: Isp (Solid Rocket) = 270 sec

Isp (Liquid Rocket) = 455 sec

3. The hypersonic second stage was assumed to be the same as the one presented by Smart and Tetlow in “Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion”.

Source: Smart, M. K., & Tetlow, M. R. (2009). Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion. Journal of Spacecraft and Rockets, 117-125.

Specific Impulse (Isp) assumption:

◦ Isp = 1100 seconds

Source: Smart, M. K., & Tetlow, M. R. (2009). Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion. Journal of Spacecraft and Rockets, 117-125.

Parameter Value

Total Mass 2000 kg

Structural Mass 1200 kg

Propellant Mass 800 kg

Propellant Liquid Hydrogen

Specific Impulse (Isp) 1100 sec

Burn Time 456 sec

Length 16.66 m

Aplanform 24.01 m2

Initial Velocity Mach 6 = 3360.35 m/s

Initial Altitude 28 km

Final Velocity Mach 12 = 4080 m/s

Final Altitude (at Burnout) 38 km

Thrust (N) 30000 N

Source: Smart, M. K., & Tetlow, M. R. (2009). Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion. Journal of Spacecraft and Rockets, 117-125.

1st Stage sizing ◦ The final conditions had to be :

altitude = 28km

burnout velocity = Mach 6

Source: Curtis, H. D. (2010). Chapter 11: Rocket Vehicle Dynamics. In H. D. Curtis, Orbital Mechanics for Engineering Students (pp. 656-688). New York: Elsevier.

Iterative procedure used to determine optimal values for the thrust and burn time of the rocket

Parameter Value

Thrust (N) 710000

Specific Impulse (Isp) (seconds) 270

Exhaust Velocity (m/s) 268.33

C (m/s) 2646

Burn Time (seconds) 22

Initial Mass (mo) (kg) 7980

Final Mass (mf) (kg) 2076.6

Tmax (seconds) 363.5

n (mass ratio) 3.84

Burnout Height (hbo) (km) 28

Burnout Velocity (Vbo) (m/s) 3350

Since the burnout flight properties are already known, the coasting flight properties needed to be calculated for the 2nd hypersonic second stage through the following equations:

Source: Curtis, H. D. (2010). Chapter 11: Rocket Vehicle Dynamics. In H. D. Curtis, Orbital Mechanics for Engineering Students (pp. 656-688). New York: Elsevier.

Resultant conditions for the launch vehicle at the end of the hypersonic second stage

Parameter Value

Initial Mass (mo) (kg) 2076.6

Final Mass (mf) (kg) 1276.6

Tmax (seconds) 535.2

n (mass ratio) 1.63

Burnout Height (h) (km) 38

Coasting Height (h) (km) 161.54

Total Height (h) (km) 200

Final Coasting Velocity (Vco) (m/s) 2427.8

The 3rd stage liquid fuelled rocket was then sized accordingly to successfully achieve a final altitude of 600 km

The same methodology was used as in the previous sizing:

Resultant conditions for the launch vehicle at the end of the 3rd stage are shown:

Parameter Value

Thrust (N) 10000

Specific Impulse Isp (seconds) 455

Burn Time (seconds) 16.3

Initial Mass (mo) (kg) 76.56

Final Mass (mf) (kg) 40

Tmax (seconds) 295.36

n (mass ratio) 1.91

Coasting Height (h) (km) 401.33

Assumption was made that the mass fraction for this new launch vehicle was the same as the mass fraction of the Dnepr 1 rocket.

Allows for an estimation of total mass, inclusive of structural mass to be made

Dnepr 1 Rocket 1st Stage 2nd Stage 3rd Stage

Mass Fraction 0.92 N/A 0.45

New Launch Vehicle – Mass of each

stage (kg)

6416.58 kg 2000 kg 81.22 kg

Final Vehicle Mass (kg) 8497.8 kg

Final Propellant Mass (kg) 7980 kg

Source: Isakowitz, S., Hopkins, J. B., & Hopkins, Jr, J. P. (2004). International Reference Guide to Space Launch Systems. American Institute of Aeronautics and Astronautics.

The final altitude is obtained by performing a summation of all the maximum heights achieved by each stage of the launch vehicle.

This final altitude correlates to the desired final altitude as illustrated in the initial design constraints

Final Altitude (km) 600.24 km

Propellant Costing

Cost per Kilogram (in terms of propellant)

Cost per Kilogram (from launch to orbit)

Comparison between Cost/kg (from launch to orbit) for our vehicle and other currently available launch vehicles

Determined the propellants for the various stages of the new launch vehicle

Calculated the mass of the propellants for each stage

Stage Propellant Mass of Propellant

(kg)

Isp (seconds)

1 ALICE 5903.255 270

2 Liquid Hydrogen 800 1100

3 LOX/LH 36.55 455

Propellant

Component

%

Composition

Mass of

Propellant

Component (kg)

Cost per kg

(USD)

Total Cost

(USD)

Ammonium

Perchlorate

70 % 4132.3 $2 $8264.56

R45-M Resin

HTPB

15% 885.5 $18.73 $16585.415

Nano-

Aluminium

15% 885.5 $9 $7969.5

• Final cost of propellant for the 1st Stage: $32,819.475

Source: http://www.alibaba.com/showroom/ammonium-perchlorate.html ; http://www.alibaba.com/showroom/nano-aluminium-powder.html ; www.alibaba.com

Propellant Mass of Propellant

(kg)

Cost per kg (USD) Total Cost (USD)

Liquid Hydrogen 800 $3.125 $2500

Source: US Department of Energy

Propellant

Component

%

Composition

Mass of

Propellant

Cost/kg (USD) Total Cost

(USD)

Oxygen 80% 29.24 $0.21 $6.14

Hydrogen 20% 7.31 $3.125 $22.661

Source: US Department of Energy ; http://www.chemicool.com/elements/oxygen.html ;

Another analysis was performed to cover the total cost per kilogram from launch to orbit, inclusive of both propellant and structural costs.

The Dnepr 1 rocket was chosen as the launch vehicle by which our launch vehicle could be scaled to in terms of estimating a structural cost.

To allow for conservative estimation the cost per launch of the Dnepr 1 rocket was assumed to be $11 million

Source: Isakowitz, S., Hopkins, J. B., & Hopkins, Jr, J. P. (2004). International Reference Guide to Space Launch Systems. American Institute of Aeronautics and Astronautics.

The process of calculating the cost of launch per kilogram for the new launch vehicle is shown below:

1st Stage 2nd Stage 3rd Stage

% of total mass 79.3% 19.7% 1%

Contribution to

total launch cost

$8.723 Million $2.167 Million $0.1122 Million

Propellant N2O4/UDMH N2O4/UDMH N2O4/UDMH

Propellant Mass 147900 kg 36740 kg 1910 kg

Propellant Cost $5,882,630 $1,461,569.33 $75,982.79

Structural Cost $2,840,370 $705,430.67 $36,217.21

Structural Cost/kg

of propellant

$19.204/kg $19.2/kg $18.962/kg

Cost for new 40kg

to LEO Launch

Vehicle

$113,366.12 $15,360 $673.06

Source: Isakowitz, S., Hopkins, J. B., & Hopkins, Jr, J. P. (2004). International Reference Guide to Space Launch Systems. American Institute of Aeronautics and Astronautics. ; Encyclopedia Astronautica

A general estimation for the total cost of the new launch vehicle to launch 40kg into orbit and an estimation for the cost / kilogram for the new launch vehicle is shown below:

Total Cost to Launch 40kg ($) $129,419.17

Cost / kg ($) $3,235.48

An estimation of the cost / kilogram to launch to orbit was calculated for a case where the hypersonic second stage was reused 5 times.

The new total cost for each stage factoring in the above case is shown below:

Cost / kilogram (launch to orbit) = $2928.28

1st Stage 2nd Stage 3rd Stage

Cost for new 40kg

to LEO Launch

Vehicle

$113,366.12 $3072 $693.06

Table compiled by Jendi Kepple

Australia has world leading capability and expertise in the field of hypersonics

Benefits of new research

Benefits of obtaining and transferring new skills

Profit from domestic reselling and international exportation

Through focusing on perfecting what Australia does so well, this project will successfully initiate the Australian space industry through innovation and expert application

Special thanks to: ◦ Dr Steve Tsitas

◦ Dr Andrew Neely

◦ Jendi Kepple