Lunar Landing and Launch Pad Construction: Concepts and Criteria
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Lunar↗
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Space Launch System Launch Windows and Day of Launch Processes
Lunar missions benefit from varying the launch azimuth as a function of launch time to allow longer launch windows with minimum performance impacts. This variable azimuth approach allows the vehicle to track the Moon’s apparent motion due to Earth’s rotation. The Space Launch System (SLS) Block 1 vehicle design requires the mission to launch into an elliptical parking orbit to provide sufficient energy to insert Orion into a Trans-Lunar Injection (TLI) orbit. The primary benefit of varying the launch azimuth, and as a result the achieved orbit inclination, allows the SLS Interim Cryogenic Propulsion Stage (ICPS) to perform its TLI burn closer to perigee and take advantage of performing a burn in a location where the burn will optimally raise apogee.
launch window↗
Space Launch System Trans Lunar Payload Delivery Capability
NASA Marshall Space Flight Center (MSFC) has successfully completed the Critical Design Review (CDR) of the heavy lift Space Launch System (SLS) and is working towards first flight of the vehicle in 2018. SLS will begin flying crewed missions with an Orion to a lunar vicinity every year after the first 2 flights starting in the early 2020's. So as early as 2021 these Orion flights will deliver ancillary payload, termed "Co-Manifested Payload", with a mass of at least 5.5 metric tons and volume up to 280 cubic meters to a cis-lunar destination. Later SLS flights have a goal of delivering as much as 10 metric tons to a cis-lunar destination. This presentation will describe the ground and flight accommodations, interfaces, and resources planned to be made available to Co-Manifested Payload providers as part of the SLS system. An additional intention is to promote a two-way dialogue between vehicle developers and potential payload users in order to most efficiently evolve required SLS capabilities to meet diverse payload requirements.
Jackman, A. L.↗
Studies of cryogenic propellant storage and handling for the lunar landing and launch facility (complex 39L)
A brief description of Complex 39L as it is currently conceived is presented. A brief discussion of lunar thermal history is then presented. From this follows a discussion of the current lunar thermal environment which will impact the design of cryogenic storage and handling facilities on the moon. Some previous studies are discussed. A conceptual design of liquid oxygen and hydrogen storage facilities is presented. The essential feature of this facility is that cryogens are to be stored in a number of small tanks which can serve as lander propellant tanks rather than as one large storage vessel. These tanks will be placed under a Fuel Inventory Tent (FIT) for shadow shielding. Methods of dealing with propellant boil-off are discussed. A base case cascade refrigeration system for requirements are such that it seems very feasible to construct a prototype boil-off recovery system in a laboratory environment.
Linsley, Jerald N.↗
Lunar Landing and Launch Pad Construction – Concepts and Criteria
When a lander vehicle launches or lands on the Moon, the rocket engine exhaust plume impinges on the surface and interacts with the regolith to create blast ejecta and associated cratering of the surface. Lunar regolith blast ejecta travels at high velocities (>2,000 m/s) for long distances (kilometers) in a vacuum environment [1] creating hazards for surrounding assets and it can also impact the bottom of the lander vehicle, risking damage to the engines, thermal insulation and sensors. Ballistic particles can possibly enter cislunar space and achieve orbit as debris, if the ejecta is sufficiently energetic. The cratering and regolith erosion can endanger the vehicle itself by affecting the soil stability under the landing gear. Landing on unpredictable terrain with varying topography, natural craters and rock hazards is also hazardous risks tipping a lander at dangerous angles in extreme conditions that may also violate maximum slope angles for subsequent launch operations. During launch (Figure 1), an overpressure pulse created by the ignition of the rocket engines can pose significant ejecta risks to the vehicle. Dust clouds raised during landing limit the efficacy of sensors and reduce visibility for the astronaut pilots, creating significant real-time risk during landing site selection by the pilot or computer navigation system. Future lunar spaceports will require mitigations to these launch and landing risks [2]. There are four main objectives of this effort:1) To establish the state of the art in LLP construction methodologies. 2) To propose criteria for trade studies of LLP concepts. 3) To publicly share the authors’ ideas for potential LLP solutions. 4) To serve as the starting point for future development of LLP technologies. Establishing the state of the art in the area of off-Earth Launch & Landing Pad concepts will base-line the work that has been completed thus far and highlight the wide span between current Technology Readiness Levels (TRLs)and operational readiness. The authors aim to communicate the need for funding in this area in the near term by illustrating that there is much work to be completed before a truly viable option exists. Setting forth criteria for trade studies of LLP concepts is important for several reasons. The first and most straightforward is to establish a framework for performing trade studies on LLP concepts. This will enable NASA to select the most promising concepts for continued development, and it will also help technology developers understand how their concepts compare with others and the priorities of development effort.
landing pad↗
Microwave Sintering: Initial Scale-Up for Lunar Landing and Launch Pad Construction
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microwave↗
Lunar landing and launch facilities (Complex 39L): Guidance systems and propellant systems
After a general, overall definition of Complex 39L during the previous two years, the 1988-89 projects were chosen to focus on more specific aspects, specifically, guidance systems and propellant systems. Six teams or subtasks were formulated: cascade refrigeration for boil-off recovery of cryogenic storage vessels; lunar ground-based radar system to track space vehicles; microwave altimeter for spacecraft; development of a computational model for the determination of lunar surface and sub-surface temperatures; lunar cryogenic facility for the storage of fuels; and lunar lander fuel inventory tent for the storage of cryogenic vessels. At the present time, a cascade refrigeration system for a cryogenic boil-off recovery system has been designed. This is to serve as a baseline system. The ground-based tracking system uses existing technology to implement a reliable tracking radar for use on the lunar surface. A prototype has been constructed. The microwave altimeter is for use on lunar landers. It makes use of the Doppler effect to measure both altitude and the vertical velocity component of the spacecraft. A prototype has been constructed. A computational model that predicts the spatial and temporal temperature profiles of the lunar subsurface was formulated. Propellant storage vessels have been designed. A support for these vessels which minimizes heat leaks was also designed. Further work on the details of the Fuel Inventory Tent (FIT) was performed. While much design work on the overall Complex 39L remains to be done, significant new work has been performed in the subject areas.
Source record↗
The Impact of Traffic Prioritization on Deep Space Network Mission Traffic
A select number of missions supported by NASA's Deep Space Network (DSN) are demanding very high data rates. For example, the Kepler Mission was launched March 7, 2009 and at that time required the highest data rate of any NASA mission, with maximum rates of 4.33 Mb/s being provided via Ka band downlinks. The James Webb Space Telescope will require a maximum 28 Mb/s science downlink data rate also using Ka band links; as of this writing the launch is scheduled for a June 2014 launch. The Lunar Reconnaissance Orbiter, launched June 18, 2009, has demonstrated data rates at 100 Mb/s at lunar-Earth distances using NASA's Near Earth Network (NEN) and K-band. As further advances are made in high data rate space telecommunications, particularly with emerging optical systems, it is expected that large surges in demand on the supporting ground systems will ensue. A performance analysis of the impact of high variance in demand has been conducted using our Multi-mission Advanced Communications Hybrid Environment for Test and Evaluation (MACHETE) simulation tool. A comparison is made regarding the incorporation of Quality of Service (QoS) mechanisms and the resulting ground-to-ground Wide Area Network (WAN) bandwidth necessary to meet latency requirements across different user missions. It is shown that substantial reduction in WAN bandwidth may be realized through QoS techniques when low data rate users with low-latency needs are mixed with high data rate users having delay-tolerant traffic.
prioritization↗
The Impact of Traffic Prioritization on Deep Space Network Mission Traffic
A select number of missions supported by NASA's Deep Space Network (DSN) are demanding very high data rates. For example, the Kepler Mission was launched March 7, 2009 and at that time required the highest data rate of any NASA mission, with maximum rates of 4.33 Mb/s being provided via Ka band downlinks. The James Webb Space Telescope will require a maximum 28 Mb/s science downlink data rate also using Ka band links; as of this writing the launch is scheduled for a June 2014 launch. The Lunar Reconnaissance Orbiter, launched June 18, 2009, has demonstrated data rates at 100 Mb/s at lunar-Earth distances using NASA's Near Earth Network (NEN) and K-band. As further advances are made in high data rate space telecommunications, particularly with emerging optical systems, it is expected that large surges in demand on the supporting ground systems will ensue. A performance analysis of the impact of high variance in demand has been conducted using our Multi-mission Advanced Communications Hybrid Environment for Test and Evaluation (MACHETE) simulation tool. A comparison is made regarding the incorporation of Quality of Service (QoS) mechanisms and the resulting ground-to-ground Wide Area Network (WAN) bandwidth necessary to meet latency requirements across different user missions. It is shown that substantial reduction in WAN bandwidth may be realized through QoS techniques when low data rate users with low-latency needs are mixed with high data rate users having delay-tolerant traffic.
quality of service↗
Lunar mission performance evaluation procedure for orbit-launched nuclear vehicles
Lunar mission performance evaluation procedure for orbit-launched vehicles with nuclear propulsion system
LUNAR FLIGHT↗
Return launch and re-entry vehicles.
Manned lunar return launch and reentry vehicles, emphasizing spacecraft and system weight in relation to mission mode and propulsion
WEIGHT FACTOR↗
Building a Lunar or Martian Launch Pad with the in situ Materials: Recent Laboratory and Field Studies
No abstract available
Hintze, Paul E.↗
NASA's Space Launch System: Launch Capability for Lunar Exploration and Transformative Science
No abstract available
Creech, Steve↗
Microwave Sintering: Initial Scale-up for Lunar Landing and Launch Pad Construction
No abstract provided
Bahr, Curtis W.↗
Launch windows for two types of orbits synchronous with the lunar period
Saturn/Centaur launch windows for orbits synchronous with lunar period
ORBIT CALCULATION↗
Use of lunar produced propellants for manned Mars missions
Manned Mars Mission departures from low lunar orbit (LLO), L2, and low Earth orbit (LEO), using oxygen or oxygen and hydrogen produced on the Lunar surface; or Phobos produced propellants; are compared to departures from LEO using Earth produced propellants. The economy of a given scheme is a function of the ratio of Earth launch to lunar launch costs per unit mass. To achieve savings on the order of 40% of total Earth launch costs for steady state operations requires the availability of both oxygen and hydrogen on the Moon and launch per unit mass costs of lunar surface to LLO in the range of 25% of Earth to LEO costs.
Gus R Babb↗
A small scale lunar launcher for early lunar material utilization
A system for the launching of lunar derived oxygen or raw materials into low lunar orbit or to L2 for transfer to low earth orbit is presented. The system described is a greatly simplified version of the conventional and sophisticated approach suggested by O'Neill using mass drivers with recirculating buckets. An electromagnetic accelerator is located on the lunar surface which launches 125 kg 'smart' containers of liquid oxygen or raw materials into a transfer orbit. Upon reaching apolune a kick motor is fired to circularize the orbit at 100 km altitude or L2. These containers are collected and their payloads transferred to a tanker OTV. The empty containers then have their kick motors refurbished and then are returned to the launcher site on the lunar surface for reuse. Initial launch capability is designed for about 500T of liquid oxygen delivered to low earth orbit per year with upgrading to higher levels, delivery of lunar soil for shielding, or raw materials for processing given the demand.
Snow, W. R.↗
Integration Of Launch Vehicle Simulation/Analysis Tools And Lunar Cargo Lander Design
Part 2, which will be discussed in this report, will discuss the development of a Lunar Cargo Lander (unmanned launch vehicle) that will transport usable payload from Trans- Lunar Injection to the moon. The Delta IV-Heavy was originally used to transport the Lunar Cargo Lander to TLI, but other launch vehicles have been studied. In order to uncover how much payload is possible to land on the moon, research was needed in order to design the sub-systems of the spacecraft. The report will discuss and compare the use of a hypergolic and cryogenic system for its main propulsion system. The guidance, navigation, control, telecommunications, thermal, propulsion, structure, mechanisms, landing gear, command, data handling, and electrical power sub-systems were designed by scaling off other flown orbiters and moon landers. Once all data was collected, an excel spreadsheet was created to accurately calculate the usable payload that will land on the moon along with detailed mass and volume estimating relations. As designed, The Lunar Cargo Lander can plant 5,400 lbm of usable payload on the moon using a hypergolic system and 7,400 lbm of usable payload on the moon using a cryogenic system.
DeJean, George Brian↗