Lunar Landing Vehicle Research Articles

Further Advances for staging orbits of manned lunar exploration mission

With the increasing demand for manned lunar exploration and deep-space exploration missions, the application value of the three-body orbit in cislunar space becomes increasingly prominent due to its special orbital characteristic. The efficient round-trip versions for manned missions of both the nominal fuel-optimal and emergency transfers between NRHO and the Moon are exploited deeply. The collocation method and indirect optimization technique with constraint gradients are utilized in conjunction with the features of the NRHO configuration to increase the algorithm's accuracy and speed. The required initial value and suitable collocation point can be determined quickly based on a global search strategy and Lambert algorithm. To highlight the effectivity of the design scheme and aim at global lunar surface exploration, typical scenarios of lunar sites at different latitudes and longitudes are investigated with various time of flight, which show that the difference of fuel consumption varies from a few kilograms to hundreds of kilograms. Meanwhile, the relations among Moon landing sites, transition time and fuel consumption are analyzed and the advantages and disadvantages of NRHO in lunar exploration are compared. The design results have some reference value for the lunar landing vehicle's deployment in the near-Moon space, as well as the round-trip transfer mission and key parameter selection.

Simulated Astronaut Kinematics and Injury Risk for Piloted Lunar Landings and Launches While Standing.

During future lunar missions, astronauts may be required to pilot vehicles while standing, and the associated kinematic and injury response is not well understood. In this study, we used human body modeling to predict unsuited astronaut kinematics and injury risk for piloted lunar launches and landings in the standing posture. Three pulses (2-5 g; 10-150 ms rise times) were applied in 10 directions (vertical; ± 10-degree offsets) for a total of 30 simulations. Across all simulations, motion envelopes were computed to quantify displacement of the astronaut's head (max 9.0 cm forward, 7.0 cm backward, 2.1 cm upward, 7.3 cm downward, 2.4 cm lateral) and arms (max 25 cm forward, 35 cm backward, 15 cm upward, 20 cm downward, 20 cm lateral). All head, neck, lumbar, and lower extremity injury metrics were within NASA's tolerance limits, except tibia compression forces (0-1543 N upper tibia; 0-1482 N lower tibia; tolerance-1350 N) and revised tibia index (0.04-0.58 upper tibia; 0.03-0.48 lower tibia; tolerance-0.43) for the 2.7 g/150 ms pulse. Pulse magnitude and duration contributed over 80% to the injury metric values, whereas loading direction contributed less than 3%. Overall, these simulations suggest piloting a lunar lander vehicle in the standing posture presents a tibia injury risk which is potentially outside NASA's acceptance limits and warrants further investigation.

Key Technologies, Systems, and Infrastructure Enabling the Commercialization and Human Settlement of Low Earth Orbit, the Moon, and Cislunar Space

More than 50 years have passed since 2001: A Space Odyssey debuted in April 1968. In the film, Dr. Heywood Floyd flies to a large artificial gravity space station orbiting Earth aboard a commercial space plane. He then embarks on a commuter flight to the Moon arriving there 25 h later. Today, in this the 52nd anniversary year of the Apollo 11 lunar landing, the images portrayed in 2001 still remain well beyond our capabilities. This article examines key technologies and systems (e.g., in situ resource utilization, fission power, advanced chemical and nuclear propulsion), and supporting orbital infrastructure (providing a propellant and cargo transfer function), that could be developed by industry for both NASA and future commercial ventures over the next 30 years, allowing the operational capabilities presented in 2001 to be achieved, although on a more Spartan scale. Lunar-derived propellants (LDPs) will be essential to developing a reusable lunar transportation system that can allow initial outposts to evolve into settlements supporting a variety of commercial activities. Deposits of icy regolith discovered at the lunar poles can supply the feedstock material needed to produce liquid oxygen (LO2) and liquid hydrogen (LH2) propellants. On the lunar nearside, near the equator, iron oxide-rich volcanic glass beads from vast pyroclastic deposits, together with mare regolith, can provide the feedstock materials to produce lunar-derived LO2 plus other important solar wind implanted (SWI) volatiles, including H2 and helium-3. Megawatt-class fission power systems will be essential for providing continuous “24/7” power to processing plants, human settlements and commercial enterprises that develop on the Moon and in orbit. Reusable lunar landing vehicles will provide cargo and passenger “orbit-to-surface” access and will also transport LDP to Space Transportation Nodes (STNs) located in lunar polar (LPO) and lunar equatorial orbits (LLO). Reusable space-based, lunar transfer vehicles (LTVs), operating between STNs in low Earth orbit (LEO), LLO, and LPO, and able to refuel with LDPs, offer unique mission capabilities, including short transit time crewed cargo transports. Even commuter flights similar to that portrayed in 2001 appear possible, allowing 1-way trip times to and from the Moon as short as 24 h. The performance of LTVs using both RL10B-2 chemical rockets and a variant of the nuclear thermal rocket (NTR), the LO2-Augmented NTR (LANTR), are examined and compared. If only 1% of the LDP obtained from icy regolith, volcanic glass, and SWI volatile deposits were available for use in lunar orbit, such a supply could support routine commuter flights to the Moon for many thousands of years. This article provides a look ahead at what might be possible in the not too distant future, quantifies the operational characteristics of key in-space and surface technologies and systems, and provides conceptual designs for the various architectural elements discussed.

Effects of rocket engines on laser during lunar landing

Abstract In the Chinese moon exploration project “ChangE-3”, the laser telemeter and lidar are important equipments on the lunar landing vehicle. A low-thrust vernier rocket engine works during the soft landing, whose plume may influence on the laser equipments. An experiment has first been accomplished to evaluate the influence of the plume on the propagation characteristics of infrared laser under the vacuum condition. Combination with our theoretical analysis has given an appropriate assessment of the plumeʼs effects on the infrared laser hence providing a valuable basis for the design of lunar landing systems.

The challenges of designing a lightweight spacecraft structure for landing on the lunar surface

The Johns Hopkins University Applied Physics Laboratory (JHU/APL) has been working with NASA's Marshall Space Flight Center (MSFC) on a lunar lander design that would take scientific measurements on the surface of the moon. This effort is part of NASA's Robotic Lunar Lander (RLL) Development Project. The requirements imposed on the design of the lander are: (1) Provide a lightweight lander structure to minimize the launch costs and maximize the payload carrying capability, (2) Minimize the lander launch envelope to allow for launching multiple landers on a single launch vehicle, (3) Given specific approach velocities, design a lander with geometric properties (low center-of-gravity, etc.) that maximizes the chances for a controlled landing on the lunar surface, (4) Provide a stable platform for all of the various scientific instruments. The lightweight lander requirement originates from the desire to minimize the launch costs and possibly package multiple landers on a single launch vehicle. The use of lightweight composite materials and advanced manufacturing techniques are employed throughout the design and construction of the structure in order to minimize mass and maximize structural stiffness. Minimizing the launch envelope enables the potential packaging of several spacecraft into one launch vehicle shroud. By having multiple landers, the scientific return is enhanced. Multiple spacecraft on the lunar surface provides independent confirmation of science measurements taken and also highlights any variance in the science data taken at differing lunar latitudes. Naturally, the launch cost per lander is greatly reduced if more than one lander can be packaged on a single launch vehicle. The lunar lander vehicle must arrive at the lunar surface at an upright orientation. In order to accomplish this, the structure geometry must be designed to accommodate attitude errors in roll, pitch and yaw. In addition, the structure must be able to withstand various landing scenarios, which include a wide range of landing velocities, decelerations and attitude rates. The stability requirement is dictated by the nature of the instrument suite. Many of the measurements to be taken on the lunar surface are so sensitive that any mechanical disturbances could potentially corrupt the measured data. Mechanical disturbances such as operating mechanisms, thermal expansion/contraction, deployable booms or other devices must be minimized. It is imperative that all of the structural requirements are met for this mission. Failure to meet any of the numerous challenging requirements could cause the loss of all the scientific objectives.

Lunar spacecraft powered descent control using higher-order sliding mode techniques

Robust performance of a flight control system in the presence of parametric uncertainty and external disturbances is of paramount importance to a successful planetary exploration program. The present research is concerned with the design of an autopilot that uses high-order sliding mode (HOSM) control principles so as to enhance the robustness properties of a lunar landing vehicle during the approach phase of powered descent. The design technique is applied to a high-fidelity simulation of the Apollo Lunar Module (LM). The design efficiently utilizes both the reaction control system (RCS) actuators and the severely rate-limited gimbal drive actuator (GDA) to effect smooth detection and compensation of sensed angular acceleration disturbances about the vehicle's control axes. The integration of a HOSM control law for the RCS effectors with a HOSM disturbance observer is shown to provide performance comparable to that of the heritage autopilot and may also avoid some difficulties encountered in the Apollo flights. Performance is maintained with the controller implemented in discrete time in the presence of a realistic vehicle and sensor model, demonstrating a unique application of sliding mode control to a complex aerospace system.

That One Small Step

This article highlights that three approaches for the Apollo mission were considered and investigated early in the program: direct ascent, Earth orbit rendezvous, and lunar orbit rendezvous. Direct ascent would entail a direct shot from Earth to the moon, requiring an enormous rocket assembly, named the Nova rocket that required 15 first stage engines and would dwarf the Saturn V eventually selected as the launch vehicle. It also required a massive lunar landing vehicle to return the astronauts from the moon directly to Earth. At liftoff, the first stage burned 15 tons of fuel a second, requiring approximately 50,000 horsepower to power the fuel pumps to feed the engines. The Apollo 13 movie followed the actual flight with a fair degree of accuracy, recognizing that it had to compress four days of real-life tension into a two-hour motion picture. The film dramatized the explosion of the oxygen tank by showing the astronauts being thrown about in the cabin. In reality, the astronauts only heard a bang and then the warning alarm for low electrical bus voltage.

Nonlinear transformation methods for gravity-turn descent

Introduction •RAVITY-TURN guidance has been widely investigated 1 ~ and used in practice' for terminal descent to a planetary surface. Previous studies of this method have utilized simple linear feedback laws to track predefined descent profiles. This Note points out that such approaches have much in common with modern nonlinear transformation methods. Furthermore, nonlinear methods can now be used for the design of descent guidance in a systematic and rigorous manner, enhancing the stability domain of the system. These matters have also been investigated for atmospheric entry by applying nonlinear methods to the Space Shuttle reference drag guidance law. Finally, a systematic new guidance law is presented that requires only slant range, slant range rate, and local vertical information. This law dispenses with the explicit need for a radar altimeter for terminal descent. Such a guidance law may have benefits for low-cost lunar landing vehicles using simple onboard sensors.

Implications of High Vacuum on the Characteristics of the Lunar Surface

The pressure on the lunar surface, which may be no higher than 10−13 terrestrial atmospheres, has already been shown to have an important effect on the behavior of postulated lunar surface materials. The observational evidence on the nature of the lunar atmosphere is summarized and the implications of the low pressure on lunar surface structural and physical properties are discussed. Lunar environmental effects including micrometeoroid impact and solar radiation are shown to be contributing to the surface characteristics. The effects of low pressures on the vesiculation of lava are used to illustrate structural features which may be present on the lunar surface. Consolidation phenomena, which are expected to produce adhesion of particles on the lunar surface, and the rôles of chemical bonds and van der Waals and electrostatic forces are discussed. The effects of the high vacuum on heat-transfer phenomena on the surface are examined and the required characteristics of postulated materials, ranging from a lightweight interconnected structure to a vesicular low-density foam, are listed. The restraints imposed on lunar exploration by the absence of the atmosphere and the expected contamination caused by lunar probes and future lunar-landing vehicles are considered.

Studies of touchdown stability for lunar landing vehicles

Touchdown stability tests for lunar landing vehicles - symmetrical & asymmetrical inelastic impacts

An FM-CW Radar for Simultaneous Three-Dimensional Velocity and Altitude Measurement

An FM-CW radar is described which simultaneously furnishes three-dimensional velocity and absolute altitude information. The velocity is measured by means of a low-frequency FM-CW Doppler technique, and the altitude information by means of phase comparison on one of the sidebands of the received spectrum. The optimum radar parameters required for the two functions are completely compatible. Use of the first-order sideband permits operation of the system down to zero feet altitude; use of the second order to very nearly zero feet. Offset reference frequency techniques and narrow-band leakage elimination techniques permit system operation through zero speed and to negative speeds. Contrary to that of conventional FM-CW altimeters, the altitude measurement technique is unaffected by amplitude noise and by frequency (Doppler) noise due to vehicle motion. The radar technique is particularly applicable to helicopters, VTOL aircraft, lunar landing vehicles and possibly also to satellite rendezvous. Actual flight test results have demonstrated the feasibility of the radar technique.

Vernier exhaust perturbations on radar and altimeter systems during a lunar landing

A PROCEDURE that has been proposed to accomplish a soft lunar landing to use a high thrust main engine stage to remove most of the approach velocity in an open loop fashion, and a low thrust vernier engine stage to control the spacecraft to a soft landing. The vernier stage employs terminal sensors to measure altitude [altimeter] and velocity [Doppler], throttling to control descent rate, and may employ body maneuvering to control horizontal velocity. It is the purpose of this paper to make an estimate of the interference that the vernier rocket exhaust may exert on the performance of the Doppler and altimeter systems of the lunar landing vehicle. The subject of rocket exhaust interference on missile guidance and communication has been under considerable study in the past. The most detailed on this topic are in the form of unpublished company reports and studies. We shall make use of the results and conclusions of some of these studies to provide upper bound estimates of flame effects on the guidance of the lunar vehicle. The main results of these are as follows: 1) Molecular absorption and polarizability are insignificant in determining the electrical properties of the rocket exhaust; the free electrons play the dominant role. 2) The electron collision frequency is determined (and therefore predictable), in the main, by the presence of those molecules with permanent electric dipole moments. 3) The observed degree of ionization of liquid rocket exhausts at exit plane conditions is enormously larger than predicted by equilibrium theory. The ionization can be explained as due to the thermal ionization of low-ionization fuel impurities (e.g., sodium) in the combustion chamber followed by slow electron-ion recombination as the combustion products accelerate, expand, and cool through the thrust chamber. The possibility of chemi-ionization as the cause of the observed ionization is not ruled out, however. 4) Computer programs were developed to predict the spatial distribution of the exhaust gases (flow pattern) for vacuum operation of the rocket motors. 5) The exhaust gas flow patterns combined with knowledge of electron collision frequency and fractional ionization (considered frozen beyond the exit plane) allow prediction of the complex electrical conductivity throughout the exhaust flare. The propagation of electromagnetic waves (of not too large

An FM-CW Radar for Simultaneous Three-Dimensional Velocity and Altitude Measurement

An FM-CW radar is described which simultaneously furnishes three-dimensional velocity and absolute altitude information. The velocity is measured by means of a low frequency FM-CW Doppler technique and the altitude information by means of phase comparison on one of the sidebands of the received spectrum. The optimum radar parameters required for the two functions are completely compatible. Use of the first order sideband permits operation of the system down to zero feet altitude. Offset reference frequency techniques and narrow band leakage elimination techniques permit system operation through zero speed and to negative speeds. Contrary to conventional FM-CW altimeters, the altitude measurement technique is unaffected by amplitude noise and by frequency (Doppler) noise due to vehicle motion. The radar technique is particularly applicable to helicopters, VTOL aircraft and lunar landing vehicles, and possibly also to satellite rendezvous. Actual flight test results have demonstrated the feasibility of the radar technique.