Presentation of Undergraduate Research Thesis

1. INTEGRATION OF A NUCLEAR THERMAL ROCKET ENGINE ONTO
A LUNAR HOPPER
Mr. Michael J. Boazzoa,b
The Ohio State University, United States, boazzo.2@osu.edu
Advisor
Dr. John M. Horacka,b
The Ohio State University College of Engineering, United States, horack.1@osu.edu
1The Ohio State University, College of Engineering, Columbus OH 43210, United States
2Battelle Center for Science, Engineering, and Public Policy at The Ohio State University, John
Glenn College of Public Affairs, Columbus OH 43210, United States

2. Overview
• Historical Recap
• Problem Statement
• Solution: Develop a Lunar Hopper
• Overview of Nuclear Thermal Propulsion
• Setting up the Optimization Problem
• Phases of Flight and How the Program Operates
• Creating a Conceptual Design
• Mass Allotments
• Individual Subsystems
• Future Work
• Appendix
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3. 1969-1972: Six Manned Landed were completed
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4. Present Day:
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“At the direction of the President of the United States, it is the stated policy of
this administration and the United States of America to return American
astronauts to the moon within the next five years,” – Vice President Mike Pence,
March 2019

5. 5

6. Eventual Goal: Permanent and
Sustainable Human Presence
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7. But where do we go, specifically?
• Despite having completed 6 manned landings, the lunar surface is still largely
unknown (5% of the surface has been explored)
• Scientists have only discovered water deposits in the past 10 years via impact probes
• Potential landing sites are mere educated guesses
• Terrain maps have been produced, but it is not enough
• If mankind is going to permanently settle the Moon, we need to pick the right spot
based on solid data observed from physical samples
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8. 8

9. 9
Developing a lunar hopper
• Here we develop the concept of a hopper: a vehicle that can take off and land
repeatedly, gather soil samples and give us as many data points as possible
• Research problem: How can we maximize the number of takeoff and landings
(hops)?
• Solution: Taking advantage of the Moon’s 1/6 gravity, we find the optimum Mass
Ratio and design a lander around a more efficient Nuclear Thermal Propulsion
System
• We use MATLAB software to set up an optimization problem to figure out what
engine, dry-mass and MR would allow us to visit the greatest number of landing sites
• We use that data outputted from MATLAB to then develop a conceptual design,
connecting theory to real life

10. Project Rover (1955-1973)
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• From 1955 to 1972, Project Rover, built
and tested flight hardware components for
NTP.
• The most famous design was the NERVA
rocket engine, which was able to output
an ISP of 900+ seconds (vacuum)
• NASA seriously considered using NTP as
a part of the Apollo Applications Program
• With budget cuts and proposed missions
to Mars cancelled, NERVA was canceled
• In 2019, NASA is currently developing next
generation NTP technologies

11. OVERVIEW OF NUCLEAR
ROCKETS
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Higher Thrust Less Thrust (but we are in the Moon’s gravity)
Less Efficient ~2/3 Times More Efficient
Can be operated in Atmosphere or Space Designed for Operations in Space Only

12. The Advantage Visualized
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13. Setting up the optimization problem
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14. Goals and Assumptions
• Primary Goal: maximize the number landing sites visited or “hops”(5 km, 10 km, 25
km)
• Find and highlight the most optimum MR (ranging from 1.5-9 in increments of 0.375)
• Lander starts each “run” on the surface with a full tank of propellant
• Uniform Lunar Gravitational Gradient (g=1.63 m/s^2)
• The Moon’s Surface is flat
• The Apolune/Apex is unconstrained
• 2-D Kinematic equations are utilized for all phases of flight (solved iteratively)
• All units are in SI
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15. Engines Selected
Critically-Limited Small Nuclear Reactor Engine (SNRE) Peewee
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16. Engine Data
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17. Algorithm Structure
• Selection of restart altitude/apolune (h_ratio) (For-Loop)
• Selection of MR (For-loop)
• Timer (For Loop)
• Ascent Function
• Coast Function
• Descent Function
• Data Extraction
• Trajectory Plotting
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18. Ascent
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19. 19
Ascent continued

20. Coast
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21. Powered Descent
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22. 22
Results

23. Critically-Limited
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• The Critically-Limited engine produced the least-satisfactory performance.
• It was viewed a primary candidate considering its low mass and low thrust that could
yield better performance.
• With the trajectories the same across the board, it completed the smallest number of
hops for all three travel distances.
• For the 5 km, 10 km and 25 km trajectories it completed a maximum of 21,10,6 hops
respectively.
• All the restart altitudes also tend to be high with relation to where the engine starts
above the ground, this is do the engine being under-powered in the lunar
environment for this specific application.

24. SNRE & Peewee
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• The SNRE with its 3500 kg dry mass was able to attain for the 5 km, 10 km, 25 km
trajectories a maximum of 28, 17 and 9 hops respectively.
• The Peewee Engine with its 4500 kg dry mass was able to attain for the 5 km, 10 km,
25 km trajectories a maximum of 30, 19 and 10 hops respectively.
• For both landing configurations, the ℎ 𝑟𝑎𝑡𝑖𝑜 typically favored values of approximately
0.4-0.45 for all three trajectory types.

25. 25
Best Hopper Combinations

26. Choosing a design
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• On paper, the Peewee/4500 kg gave a better performance by a close margin
• The Peewee is much larger by volume and heavier than the SNRE
• This would likely add to cost and complexity
• Dry mass is more likely to increase based on needed structural support
• Also must apply volume constraints (new upcoming commercial launch vehicles can
launch with 6-meter fairings)
• Our goal is to minimize the total mass so we can travel to the moon in one-trip using a
direct trans-lunar injection burn
• Keeping in mind the number of ground-breaking technologies that need to be
integrated, it makes sense to start-off small
• SNRE/3500 kg is selected as the final design

27. Trajectory Plots
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28. Refining the Trajectory Data
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29. Vehicle Design
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30. Design Graphic
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31. Ballooned Drawing
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32. Dimensioned Drawing
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33. Mass Budget
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34. Fuel Tank
• Two different materials were selected for analysis: Stainless Steel 301 and Carbon Composites.
• For this study, an existing space vehicle that utilized LH2 was selected as a starting point.
• The Centaur upper stage was selected for its historical use of a LOX/LH2 booster that was specifically designed for in-
space (non-atmospheric) operations.
• According to original NASA Technical Memorandums, the booster utilized a thickness of approximately 0.014 inches for
most of the tank (NASA TM X-1844, 1970).
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35. Octagonal Base
• The landing chassis was designed primarily to support the LH2 fuel tank, engine
mount and landing legs
• Many lessons learned possible were borrowed from the Apollo LM since the scale
is similar in terms of mass
• Aluminum 7075 was chosen as the material thanks to its lightweight characteristic
while also having enough tensile strength to support the weight of the fuel tank
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36. Landing Legs
• The mass of the landing gear was approximated to be 2.5% of the total mass of the
hopper based on data from the Apollo LM
• The Apollo LM and other unmanned landers such as a Viking have only utilized and
designed landing gear for one-time use
• Likely the final critical design of the lander gear will feature some sort of shock
absorber: metal bellows, electromagnetic and electromechanical
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37. Engine
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38. Controls
• With the philosophy of minimizing mass kept in the forefront of the design, RCS and
related propellants were seen to be not satisfactory for this application
• The Honeywell M160 was selected with each model weighing 44 kg and consuming
217 watts of power (Leve, 2009).
• Two CMGs will be utilized in order to ensure redundancy if one of them fails
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39. Flight Computers and GNC
• NASA programs such as Project ALHAT and COBALT have developed and tested
the next generation of GNC algorithms for lander vehicles
• Both programs have focused on integrating a hazard detection system composed
of lidar doppler radar, lasers and sensor technology for the vehicle to autonomously
land safely with pinpoint precision
• The final GNC and Flight Software products that came from ALHAT and COLBALT
would be a perfect match for the hopper
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40. Power
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41. Mass from Specific Power
Systems
(Larson & Wertz, 1997, p.334)
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42. Final Vehicle Performance
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43. FUTURE WORK
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Traveling 59.8 km/s
RENDEVOUS in 13.7 years
550 AU in 36 Years
NASA SLS BLOCK 1
• Work on making turbopump engines more “restart” friendly
• Integrate design studies on Zero-Boil Systems to conserve propellant
• Study on radiation from the reactor
• Do more advanced analysis on Structural Stress with Finite Element Method
• Figure out adequate method for Thermal Shielding
• Design tools for scooping up and analyzing soil properties of the lunar dust

44. THANK YOU TO ALL WHO MADE
THIS POSSIBLE !
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Mechanical and Aerospace Engineering
Dr. John Horack
Dr. Elizabeth Newton
Dennis Scott
Andrew Steen
Taylor Watson
Luke McNamara
Ron Sostaric
Eric Hurlbert
John Scott
Dr. Ali Jhemi
Dr. Prasad Mokashi

45. APPENDIX
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46. Ascent Equations
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47. Coast Equations
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48. Descent Equations
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49. Equations for CMGs
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50. Verifying the Structure of the Octagonal-Base
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