40 kg to LEO: A Low Cost Launcher for Australia. By Nicholas Jamieson
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1 40 kg to LEO: A Low Cost Launcher for Australia By Nicholas Jamieson
2 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)
3 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
4 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
5 Source: Heyman, J. (2009, October). FOCUS: CubeSats - A Costing + Pricing Challenge SatMagazine.
6 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.
7 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
8 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.
9 Capability Operability Reliability Economical Commercially Orientated Design Source: Nichols, E. E. (1997). The Space Launch Payload Process. The AIAA Journal,
10 Initial design constraints Assumptions made along the way Mathematical Methodology Resultant Sizing Parameters
11 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
12 Source: Google Maps
13 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,
14 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,
15 Parameter Total Mass Structural Mass Propellant Mass Propellant Specific Impulse (Isp) Value 2000 kg 1200 kg 800 kg Liquid Hydrogen 1100 sec Burn Time 456 sec Length m A planform m 2 Initial Velocity Mach 6 = m/s Initial Altitude 28 km Final Velocity Mach 12 = 4080 m/s Source: Smart, M. K., & Tetlow, M. R. (2009). Orbital Delivery of Small Payloads Using Hypersonic Airbreathing Propulsion. Journal of Spacecraft and Rockets, Final Altitude (at Burnout) Thrust (N) 38 km N
16 1 st 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 ). New York: Elsevier.
17 Iterative procedure used to determine optimal values for the thrust and burn time of the rocket Parameter Value Thrust (N) Specific Impulse (Isp) (seconds) 270 Exhaust Velocity (m/s) C (m/s) 2646 Burn Time (seconds) 22 Initial Mass (m o ) (kg) 7980 Final Mass (m f ) (kg) T max (seconds) n (mass ratio) 3.84 Burnout Height (h bo ) (km) 28 Burnout Velocity (V bo ) (m/s) 3350
18 Since the burnout flight properties are already known, the coasting flight properties needed to be calculated for the 2 nd 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 ). New York: Elsevier.
19 Resultant conditions for the launch vehicle at the end of the hypersonic second stage Parameter Value Initial Mass (m o ) (kg) Final Mass (m f ) (kg) T max (seconds) n (mass ratio) 1.63 Burnout Height (h) (km) 38 Coasting Height (h) (km) Total Height (h) (km) 200 Final Coasting Velocity (V co ) (m/s)
20 The 3 rd 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:
21 Resultant conditions for the launch vehicle at the end of the 3 rd stage are shown: Parameter Value Thrust (N) Specific Impulse Isp (seconds) 455 Burn Time (seconds) 16.3 Initial Mass (m o ) (kg) Final Mass (m f ) (kg) 40 T max (seconds) n (mass ratio) 1.91 Coasting Height (h) (km)
22 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 1 st Stage 2 nd Stage 3 rd Stage Mass Fraction 0.92 N/A 0.45 New Launch Vehicle Mass of each kg 2000 kg kg stage (kg) Final Vehicle Mass (kg) 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.
23 The final altitude is obtained by performing a summation of all the maximum heights achieved by each stage of the launch vehicle. Final Altitude (km) km This final altitude correlates to the desired final altitude as illustrated in the initial design constraints
24 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
25 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 Isp (seconds) (kg) 1 ALICE Liquid Hydrogen LOX/LH
26 Propellant % Mass of Cost per kg Total Cost Component Composition Propellant (USD) (USD) Component (kg) Ammonium 70 % $2 $ Perchlorate R45-M Resin 15% $18.73 $ HTPB Nano- 15% $9 $ Aluminium Final cost of propellant for the 1 st Stage: $32, Source: ; ;
27 Propellant Mass of Propellant Cost per kg (USD) Total Cost (USD) (kg) Liquid Hydrogen 800 $3.125 $2500 Source: US Department of Energy
28 Propellant % Mass of Cost/kg (USD) Total Cost Component Composition Propellant (USD) Oxygen 80% $0.21 $6.14 Hydrogen 20% 7.31 $3.125 $ Source: US Department of Energy ; ;
29
30 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.
31 The process of calculating the cost of launch per kilogram for the new launch vehicle is shown below: 1 st Stage 2 nd Stage 3 rd Stage % of total mass 79.3% 19.7% 1% Contribution to $8.723 Million $2.167 Million $ Million total launch cost Propellant N 2 O 4 /UDMH N 2 O 4 /UDMH N 2 O 4 /UDMH Propellant Mass kg kg 1910 kg Propellant Cost $5,882,630 $1,461, $75, Structural Cost $2,840,370 $705, $36, Structural Cost/kg $19.204/kg $19.2/kg $18.962/kg of propellant Cost for new 40kg $113, $15,360 $ to LEO Launch Vehicle 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
32 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 ($) $129, Cost / kg ($) $3,235.48
33 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: 1 st Stage 2 nd Stage 3 rd Stage Cost for new 40kg $113, $3072 $ to LEO Launch Vehicle Cost / kilogram (launch to orbit) = $
34 Table compiled by Jendi Kepple
35 Australia has world leading capability and expertise in the field of hypersonics
36 Benefits of new research Benefits of obtaining and transferring new skills Profit from domestic reselling and international exportation
37 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
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