Design of the subsystems for this lunar landing spacecraft has been divided among various member educational institutions in the Texas Space Grant Consortium (TSGC). The design team at Texas A&M University has chosen to lead the design and analysis effort for the structural components of the spacecraft. To facilitate the design, a need statement was developed. This need statement was broadened and analyzed to produce various requirements and constraints on the design. After further investigation of these needs, constraints, and technologies, several possible designs were developed and a single spacecraft and payload configuration was selected. The functional and performance requirements, design alternatives, and selected structural design are discussed at the following sections.
For this mission to be a success, the structure of the lander must maintain its structural integrity while experiencing maximum loading conditions in all mission phases. The three primary types of loading are static, dynamic, and thermal loads
Static Loads - The primary static loads arise from the average thrust loads delivered by the Delta-II launch vehicle and XLR-132 main engine. The other static loads applied to the structure will simply be the 1g load in pre-flight handling and the 1 lunar g exerted on the structure after touchdown.
Dynamic Loads - There are many factors that contribute to dynamic loads in a structure of this type. Vibration loads, shock, and acoustic noise can all adversely affect structural integrity and vehicle reliability.
Vibration loads can arise from many different sources. The primary sources of vibration excitation are acoustic pressures generated by rocket-engine operation, aerodynamic pressures created by boundary-layer fluctuations, mechanically induced vibration from rocket-engine pulsation, which is transmitted through the vehicle structure, and localized machinery. These loads are random forcing functions but have statistically predictable characteristics. These characteristics are summarized by a power spectral density plot that is unique for each launch vehicle and is provided by the manufacturer of the launch platform. The spacecraft structure must be tailored to avoid the launch vehicle's natural frequencies in order to prevent a catastrophic failure. Typical resonance sources to avoid include oscillations in the propulsion system (pogo), aerodynamic buffeting during ascent, and bending of the solid rocket motors.
Shock loads can arise from any transient event such as pyrotechnic separation (staging), firing an RCS jet, or landing on the lunar surface (impact). Pyrotechnic shock causes high acceleration and high frequency over a very short time. Because shock loads attenuate quickly, they seldom damage structures removed from the immediate impulse, but they may seriously harm nearby electronic components.
Acoustic noise is generated by turbulent exhaust flow from the propulsion system. This high-velocity flow produces pressure fluctuations, referred to as noise, and has adverse effects on the vehicle and its operations. Structures with high surface area and low mass, including skin sections and solar array panels, respond strongly to acoustic noise. Acoustic levels for the Delta-II are empirically derived from past flight data since they are difficult to predict with any reliability.
Thermal Loads - Thermal isolation of electronics and payload must be provided along with thermal control of the storable propellant lines to prevent freezing. Space vehicles commonly experience large thermal gradients because of the time-related and/or attitude-related exposure to solar radiation. The thermal effects of stress and deflections of the structure require that temperature distributions be consistent with the strength and deflection requirements. The significance and sensitivity of thermal effects are greatly reduced by the use of composite materials for the structure. These materials can be layered to achieve, within limits, the desired thermal coefficient of expansion and thereby reduce the thermal effects.
The critical factors that drive the design of the spacecraft are a function of the mission phases, as described below:
Pre-launch - Loads can be imparted to the structure prior to launch due to handling during stacking sequence and pre-flight checks. Effects of ambient temperature, humidity and oxidation on the structure and internal hardware must also be considered.
Launch and Ascent - The largest steady-state booster accelerations and acoustic noise levels occur during launch, particularly in the transonic phase (Mach 0.9 to 1.2). Transient loads also arise from propulsion system engine vibrations, stage separations, vehicle maneuvers, propellant slosh, and payload fairing separation.
Orbit and Lunar Transit - Steady-state thruster accelerations, transient loads during pointing maneuvers and attitude control burns, pyrotechnic shock from separation events, and vibrations due to antenna and solar panel deployments all contribute to the loading environment while in orbit and in lunar transit. In addition to these structural loads and thermal loads, several other factors should be considered when designing for this phase of the mission. Since the spacecraft is in microgravity, there are no natural convection effects. Also, in the vicinity of earth, micrometeroids travel with velocities on the order of 10 to 30 km/s, which makes even millimeter-sized particles potentially damaging to space structures.
The vacuum of space may cause the evaporation of a material, or a volatile component of the material in a phenomenon known as "outgassing." Although the evaporation of a component of a material may not reduce the effectiveness of the material, the deposition of the vapor on a colder surface may be intolerable, particularly for solar panels, optical components, or sensitive electronics. Metals do not usually evaporate in space at modest temperatures, but organic materials, including elastomers, plastics, coatings, adhesives, and lubricants, must be of very high molecular weight to avoid evaporation. Moving electrical-contact surfaces and bearings require either reliable isolation from the vacuum environment or special selection of materials, especially where long-time operation is involved.
Radiation is also a significant factor affecting spacecraft design. Metals and ionic compounds are relatively resistant to solar and cosmic radiation. However, semiconductors and other electrical materials are sensitive to permanent radiation damage. Organic materials are also susceptible to degradation by both electromagnetic and particle radiation, especially in vacuum. Organic polymers of high molecular weight may have such low vapor pressures that their evaporation in vacuum at reasonable temperatures is not significant. However, intense radiation can yield fragments of reduced molecular weight and increased vapor pressure that outgassing can become an issue. Radiation and magnetic field protection for sensitive internal hardware can often be provided through the use of non-load bearing skin panels or membranes.
Lunar Landing - The primary loads applied on landing are an initial shock load (impact) followed by free vibration of the structure. The possibility of contamination of the lunar soil sample by landing exhausts must also be considered.
Post-landing - After landing, the structure will operate as a platform for the oxygen production experiment. Secure platforms for communication must also be provided. Any movable mechanical systems must be sealed properly to avoid contamination by lunar debris. The system must also be robust enough to withstand the thermal environment of on lunar night (about 14 days).
In order to maximize design reliability and to minimize cost, off-the-shelf hardware should be used as much as possible to minimize cost. Composites might be better at meeting structural performance requirements but the weight savings may not be worth the cost increase. If aluminum will still meet the weight and strength requirement, it should be used. High-tech materials and complex designs should not be implemented for their own sake.
|Maximum Weight -||75 kg (structure)|
|520 kg (total vehicle weight including propellant)|
|Source: Delta-II payload requirements and initial mass breakdown|
|Maximum Volume -||Propellant Tanks - 64.1 cm diameter sphere (4 times)|
|Source: Preliminary delta-V calculations|
|Vehicle Volume - 254 cm diameter by 203 cm height (cylindrical)|
|254 cm diameter by 290 cm height (conical)|
|Source: Delta-II payload bay geometry|
|Engine Thrust Loads -||Main Engine Maximum Thrust: 12,000 N|
|Source: Rocketdyne XLR-132 main engine specification|
|3 Auxiliary Landing Engines: 500 N each|
|Source: UT - Structures and Systems Integ. Final Design Review|
|Propellant -||245 kg of Monomethyl Hydrazine (MMH) and Nitrogen Tetroxide (N2O4)|
|Source: UT - Structures and Systems Integ. Final Design Review|
|Acceleration Loads -||7.7 g's axial (steady state)|
|4.0 g's axial (dynamic)|
|2.0 g's lateral (steady state)|
|3.0 g's lateral (dynamic)|
|Source: Space Mission Design and Analysis by Wertz|
|Vibration andAcoustic Loads -||Maximum Acoustic Load: 139.6 dB (1/3 octave)|
|Minimum Lateral Frequency: 15 Hz|
|Minimum Longitudinal Frequency: 35 Hz|
|Maximum Flight Shock: 4100 g @ 1500 Hz|
|Source: Guide to Space Launch Systems|
|Thermal Loads -||Operating Temperature: -45 to +65 degrees C|
|Source: Oct. 6 Progress Report from Prairie View A&M|
|Micrometeroids -||Micrometeroid Flux: 0.02 m2/yr (1 mm diameter)|
|Velocity: 10 to 30 km/sec|
|Source: Space Vehicle Design by Griffin and French|
|Power -||Total Solar Cell Area: 100 W/m2|
|Source: Online Power Specification Sheet|
Given the design requirements, criteria, loads, and operating environment described above, we can now begin formulating conceptual designs for the structure. The load distribution or load paths in the structure depend upon the characteristics and configuration of the spacecraft and the characteristics of the structure are largely dictated by the loads. This functional dependency between loads and structures requires an iterative design process.
Any unmanned spacecraft typically consists of three primary elements: a payload, a spacecraft bus, and a booster adapter. The payload is the mission-specific equipment and its instruments (in this case the lunar oxygen plant and the rover). The spacecraft bus carries the payload and provides housekeeping functions and the booster adapter provides the load-carrying interface with the launch vehicle. The structural subsystem in the spacecraft bus provides support and alignment for the spacecraft equipment. It also cages and protects folded components during boost and deploys them in orbit.
The payload is the single most significant driver of the spacecraft design. Its physical parameters (size, weight, and power) dominate the physical parameters of the spacecraft. In the initial stages of spacecraft design the dry weight of the spacecraft was taken as some percentage of the payload. This facilitated an initial estimate of the propellant requirements given the delta-V's required for mission operation. Since propellant is usually a very large percent of the total vehicle weight and volume, the propulsion and propellant requirements were determined early.
There were many factors considered when determining the layout of each vehicle design alternative. The first two systems that were located for each design alternative were the payload and the propulsion system for the reasons discussed above. The propellant tanks are located as close to the vehicle's center of gravity (CG) as possible to minimize CG movement as propellant is burned. Translational RCS jets are aligned through CG to prevent the addition of extraneous torques to the vehicle's motion. Rotational jets are mounted in opposing pairs located far from the CG to produce a pure couple moment on the spacecraft for attitude control. Care had to be taken when locating the RCS thrusters to prevent them from contaminating sensors, antennas, and solar array cells with propellant exhaust gases. The propulsion system must be supported in such a way that the thrust vector remains aligned with the center of mass. Since the propulsion system has significant weight, it is most desirable for it to placed at the base of the spacecraft stack, near the launch vehicle interface, to help minimize structural weight. Not all of the conceptual designs considered followed this rule-of-thumb however.
The inertial measurement unit (IMU) is also located as near to the CG as possible to minimize errors in the IMU's measurements due to vehicle flexure between the CG and the IMU's location. The IMU is mounted on a rigid platform to maximize the accuracy of the measurements. The computers and other electronics are kept as near the center as possible to help shield them from possible radiation upset events and to maximize fault tolerance. All electronics are mounted on a cold plate that circulates coolant fluid and maintains the proper operating temperature.
The thermal control and power systems are located in various locations depending upon the design. As long as the CG remains roughly axial and low to maximize spacecraft stability on descent, their placement is not as critical. Batteries should be accessible, however, for pre-launch testing or replacement and should be placed where they will be at their optimum temperature.
The communications antenna is located on the top of the spacecraft since that will maximize the coverage and ensure that a hemispherical field of view is available during times of signal transmission or reception. If this location proves to be a problem, it could also be mounted on an appendage that is stowed during launch and deployed on orbit for an unobstructed view. Sensing devices (such as star trackers and horizon sensors) also require specific fields of view and pointing clearances.
Since the operation of the oxygen production plant experiment is the primary design driver for this spacecraft, four different vehicle configurations were considered with the plant operating in different ways. These different conceptual designs are described below:
Option 2 keeps the oxygen plant onboard and retrieves the sample from the rover with the use of a base deployed sample retrieval system. This option has a smaller payload area requirement in the base since only the rover and the sample retrieval container are deployed through the base. This results in some improvement in the weight and load path efficiency of the structure although the structure still must still be arranged around the rover in the base of the spacecraft. Rover deployment is still a simple process but the sample insertion is more complex. The rover arm must now place the sample into the sample retrieval container and the retrieval system must then retract and place the sample into the oxygen plant.
Option 3 eliminates the onboard sample retrieval system altogether and relies on the robotic arm on the rover for sample insertion into the oxygen plant. This option requires that the oxygen plant be located in the base of the spacecraft and that the robotic arm on the rover have a very large reach and degree of articulation. The volume constraints are as stringent as those in design Option 1 since all of the payload is located in the base of the lander. The load path and structural weight is also very inefficient for the same reasons.
Option 4 has all payload located in the top of the spacecraft. This allows for a much more efficient design of the thrust structure and for the rest of the spacecraft in general. The thrust load paths are direct and do not have to be routed around a payload in the base. Since both the oxygen plant and the rover are deployed from above the spacecraft components to the lunar surface, operation of the experiment is simplified at the cost of a more complicated method of deployment. If the oxygen plant must maintain a hard wiring link to the lander for power supply and data transmission, the side-mounted actuator deployed platform concept will probably be implemented. If the oxygen plant has onboard batteries and the experimental data can be remotely relayed to the lander for transmission to Earth, the ramp deployment concept may be employed.
The table below summarizes the relative strengths and weaknesses of the four design options.
|Option 1||Option 2||Option 3||Option 4|
|Load Path||indirect and inefficient||more efficient||indirect and inefficient||direct, most efficient design|
Figure 6 - Bottom View
Figure 7 - Bottom View (Close Up)
Figure 8 - Spacecraft in Delta-II Payload Bay
The primary goals of the structures team in the project are listed as follows: