1. The Novel Design of a Reusable Orbital Transfer Vehicle for
Payload Supply Missions to the Earth-Moon Lagrange Point Two
The IMPACT Senior Design Team: Andrew Hull, Bryan Johnson, John Kelley, Sun Jae Kim, Julie Orbin, Lucas Porter
Faculty Advisor: Dr. Javid Bayandor Manager: Melina De La Hunt
Acknowledgements: NASA Engineering and Systems Branch - Langley Research Center, Josh Korsness, Kian Sharafi, Liza Kossobokova, Ravi Saripella, Steven Chung
Background and Objectives
The objective of the IMPACT team is to create a detailed mission and vehicle
design for a reusable orbital transfer vehicle. The objective of our mission is to
design the vehicle to transport at least a 50,000 lb payload from a propellant
depot in lower earth orbit to a gateway space station located at Earth-Moon
Lagrange points one or two. The
eventual goal of the orbital transfer
vehicle would be to support Lunar,
Mars, or near earth asteroid
missions over a period of five
years or ten round trip transport
missions.
Mission Design and Analysis
Propulsion System
Launch and Assembly
Falcon Heavy
53,000 kg to LEO
$85 mil per launch
Five total Launches
Propulsion
Propulsion
Payload
Exoskeleton
Manned Module
Manned Module
LIFE SUPPORT SYSTEM
Proper recycling of critical resources such as oxygen and water
will reduce mass requirements for each mission by 240 kg,
a 4% mass savings and $60K cost savings
PRESSURE VESSEL DESIGN
Pressurized to 1 atm with 20% O2, 80% N2
Thickness set at 3 cm Using Aluminum alloy
THERMAL REGULATION
15 kW of heat from onboard
Instruments is dissipated using a
heat exchanger loop utilizing water
and ammonia for the working fluid.
Worst case scenario at LEO orbit: 394 K
Resulting surface area of radiator: 43 m2
LEO | EML2
EML2 | LEO
A trade study found that
the OTV needs to perform a
direct transfer from LEO to
EML2 with a total change
in velocity of 3.938 km/s
The true anomaly was
varied from 120 to 180
degrees in order to find
the optimal transfer orbit
A unique reusable staging system
was designed to reduce propellant
mass. It was found that staging
at 65.63% of the way through the
initial burn leads to propellant
savings of nearly 32,000 kg
A symplectic Euler method was used to map the trajectories of
the first stage module. To reuse the propulsion module and dock
with the propellant depot, a phasing and docking maneuver are
necessary. The resulting apogee of the elliptical orbit is 27,000
km and the flight time of the first stage module is 7.7 hours
1st Stage2nd Stage
Four RL60 Engines
built by Pratt &
Whitney
Specifications:
Isp: 465 s
Thrust: 250 kN
Mass: 400 kg
Exoskeleton Analysis
Fatigue analysis found that the structure
withstands up to 9.9 million cycles during
its mission
Payload
Three Part Exoskeleton
Main Structure
Two Bay Doors for Entry
and Removal of Payload
Water Vapor
Liquid Water
Ammonia Vapor
Liquid Ammonia
Cold
Plate
Internal External
Stress analysis found that the structure
withstands the 736 kN thrust generated
by the four RL60 engines chosen
Results
The developed OTV can transfer to and from LEO and EML2 in 4.5 days and
is able to support four astronauts for up to 52 days and function autonomously.
The payload bay has a capacity of more than 50,000 lbs
that is propelled by a two stage propulsion system using
bipropellant liquid hydrogen/liquid oxygen rocket
engines. The OTV is structurally capable
to withstand the number or thrust
forces and number of engine
cycles over10 missions
spanning 5 years.
OTV Transfer Orbit
First Stage Orbit
Phasing Orbit
Propellant Depot