Structures CDR - A&M
Structures Subsystem - Conceptual Design Report
Texas A&M University


Since the Apollo missions and humans' visit to the moon, only a small fraction of the Moon's surface has been explored and only limited samples from those explored areas were returned to Earth. Much remains to be learned about the moon and its potential uses in the future. From the early remotely operated lunar vehicles to the manned Apollo missions, it is known that the moon contains many resources that could be used for future explorations. A primary example of this is the lunar regalith (soil). It has been revealed experimentally that this type of soil contains significant amount of oxygen which can be extracted by several different processes. Oxygen produced could be used for Earth-return fuel oxidizer or for life support. Being able to produce the needed resources on the moon would also reduce the amount of mass that would have to be launched from Earth. The primary focus of this Discovery-class mission is on the discovery and use of lunar oxygen by sending a small, lunar oxygen demonstration device to the surface of the moon. This mission is a response to the Discovery Mission Announcement of Opportunity issued by NASA headquarters on July 28, 1994.

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.

Need Statement

To provide all primary (load bearing) and secondary (component mounting) structural support to withstand the static, dynamic, and thermoelastic loads that are applied to the lunar lander spacecraft during launch, trans-lunar flight, lunar landing, and lunar surface operations.

Need Analysis

The main function of this mission is to perform regalith oxygen extraction experiments by gathering samples with a rover and reacting these samples in a catalyst chamber with hydrogen at high temperature. The spacecraft is launched on a Delta-II launch vehicle and placed into an elliptical Earth orbit. The spacecraft main engine then performs the trans-lunar injection burn and sends the spacecraft on an intercept trajectory with the moon. The main engine is also used to perform a lunar orbit insertion and circularization burn. The spacecraft deorbits and lands on the lunar surface and performs a post-landing systems check. The oxygen production plant and a sample gathering rover are deployed to the lunar surface. Samples are gathered by the rover and placed into the oxygen plant and the experimental results are relayed through the lander's communication system to the Deep Space Network on Earth.

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.

Function Structure

The following outline details the functional requirements for the spacecraft structure.

  1. Provide structural support and withstand loads
    1. Engine thrust loads
    2. Acceleration loads
    3. Vibration and acoustic loads
    4. Thermal expansion loads
    5. Internal pressure gradients

  2. Provide environmental protection for subsystems and payload
    1. Oxidation protection
    2. Humidity protection
    3. Vibration and acoustic protection
    4. Radiation protection
    5. Micrometeorite protection
    6. Lunar dust protection
    7. Outgassing protection
    8. Thermal protection

  3. Provide interfaces with all subsystems
    1. Lander GNC
      1. Provide mounting, support, and volume for effectors
      2. Provide mounting, support, and volume for sensors
      3. Provide mounting, support, and volume for computers and/or controllers

    2. Electric power
      1. Provide mounting, support, and volume for electric generators and batteries
      2. Provide mounting, support, and volume for solar arrays
      3. Allow for actuation of solar arrays

    3. Communication
      1. Provide mounting, support, and volume for antenna
      2. Provide mounting, support, and volume for communications electronics
      3. Allow for actuation of antenna
      4. Avoid blocking antenna's field of view

    4. Propulsion
      1. Provide mounting and separation from Delta-II launch vehicle
      2. Provide mounting, support, and volume for fuel tanks
      3. Provide mounting, support, and volume for main engine
      4. Provide mounting, support, and volume for plumbing

    5. Landing Gear
      1. Provide mounting and support for landing gear
      2. Allow for actuation of landing gear

    6. Payload
      1. Provide mounting, support, and volume for all payloads
      2. Provide deployment mechanism for moon rover

    7. Science package
      1. Provide mounting, support and volume for moon rover
      2. Provide mounting, support, and volume for oxygen plant
      3. Provide mounting, support, and volume for robotic arm

Performance Requirements

The following is a summary of the preliminary performance requirements for the structure:

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


The following assumptions are being made to initiate the design process:

Conceptual Designs

One of the early decisions that must be made in the design of any spacecraft is the design and development philosophy to be adopted, i.e. whether to utilize a modular or a payload specific design approach. Modular spacecraft provide a common interface for different payloads and facilitate many types of missions with the same spacecraft. While this modularity minimizes the subsequent payload integration for later missions and subsequent operations cost, it is offset by an increased initial development cost and schedule due to the design of a more generic payload interface for many different payload requirements. A payload specific design approach centers around a single payload for one time use. The spacecraft is basically built around the payload, allowing the payload specific structural support to double as primary structure for the spacecraft. This type of concurrent engineering allows for a more weight and volume optimized spacecraft as well as a reduced development cost and schedule. The major drawback with this approach is that an entirely new spacecraft will most likely have to be designed for any payload that is significantly different from the payload the original spacecraft was designed to support. Since Discovery class missions have a fixed-cost requirement for development, the payload specific design approach was selected since it will minimize the launch weight as well as design and development costs, with the realization that the spacecraft will be utilized for this payload only.

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:

Design Option 1 - Base Deployed Oxygen Plant and Rover
In this design, the oxygen plant and rover are both placed onto the lunar surface with a common platform that is deployed through the base of the lander. The rover collects samples and places them directly into the oxygen production plant on the lunar surface. The rover and the oxygen production plant are mounted in the base of the vehicle near the main engine and thrusters until the vehicle touches down and performs a post-landing systems check. The common platform in the base is deployed with a pair of electromechanical actuators that extend vertically downward to the lunar surface. The rover disengages from the platform and moves off to extract regalith samples while the oxygen plant remains on the platform below the lander. Most all of the other spacecraft subsystems (power, thermal, GNC, communications, propellant tanks) are located above the rover and oxygen plant in the spacecraft bus. This configuration is shown in the figure below.

Figure 1: Option 1 - Base Deployed Oxygen Plant and Rover
Design Option 2 - Onboard Oxygen Plant with Base Deployed Sample Retrieval System
In this option, only the rover is deployed to the lunar surface while the oxygen production plant remains onboard the lander. The rover is deployed through the base on a smaller version of the platform deployment system described in Option 1 (since the oxygen plant is not being deployed). The rover collects the soil samples and places them into a sample retrieval container. This container is deployed through the base of the lander with a small electromechanical actuator that then brings the sample up into the oxygen plant for the experiment. The rover is mounted in the base of the lander prior to deployment while the oxygen plant is mounted at the top of the lander as shown in the figure below.

Figure 2: Option 2 - Onboard Oxygen Plant with Base Deployed Sample Retrieval System

Design Option 3 - Onboard Oxygen Plant with Direct Soil Sample Insertion
In this option, the rover is again the only part of the payload to be deployed to the lunar surface. The rover collects the samples and the robotic arm places them directly into the oxygen plant without the aid of a sample retrieval arm as in Option 2. The rover is mounted in the base of the vehicle prior to deployment. The oxygen production plant is mounted lower in the lander than in Option 2 to minimize the amount of vertical reach required by the rover's robotic arm to insert the regalith. This is shown in Figure 3 below.

Figure 3: Option 3 - Onboard Oxygen Plant with Direct Soil Sample Insertion

Design Option 4 - Side Deployed Oxygen Plant and Rover
In this last design option, both the oxygen plant and the rover are deployed to the lunar surface as in Option 1. However, rather than deploying through the base, the rover and oxygen plant are deployed from the side of the lander. This can be done with either an extendable ramp that the rover descends, pulling the oxygen plant behind it, or with a variation of the actuator deployed platform concept. In this configuration, both the rover and the lander are mounted in the top of the spacecraft prior to deployment. All subsystems (with the possible exception of the communications antenna) are mounted below the payload in the base of the lander near the engines. This is shown in Figure 4 below (ramp concept pictured).

Figure 4: Option 4 - Side Deployed Oxygen Plant and Rover

Design Comparison

The primary design drivers for the structure of this spacecraft are the location of the payload and spacecraft's subsystems. Many factors must be considered when trading off one design configuration against another, including weight, volume, efficiency of load path, ease of payload deployment, ease of operation, and cost. Payload deployment in Option 1 is a very simple process since it utilizes a base deployed payload palette. Since both the oxygen plant and the rover are on the lunar on the surface, the rover has little difficulty in placing the regalith sample into the oxygen plant for oxygen production. However, a base deployed rover and oxygen plant places severe volume constraints on both the payload and other spacecraft subsystems. An initial placement of components into a 3-D solid model of the spacecraft showed that 56 cm x 45 cm x 35 cm was available for the oxygen plant and only 50 cm x 30 cm x 25 cm for the rover. This design also requires that most other subsystems be located above the payload. Since a large amount of open area is required for base deployment, the primary structure transmitting thrust from the main engine and the thrusters must be arranged around the payload, leading to a very inefficient thrust structure and thrust load path. This consequently requires that these primary structural components be "beefed-up" in order to meet strength requirements, resulting in increased structural weight. Wiring and propellant plumbing will have to be routed around the payload in base of the spacecraft, which could prove to be difficult.

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.

Table 1 - Design Comparison

Option 1Option 2Option 3Option 4
Weight Efficiencypoormoderatepoorgood
Volume Efficiencypoormoderatepoorgood
Load Pathindirect and inefficientmore efficientindirect and inefficientdirect, most efficient design
Deploymentsimplesimplesimplemoderately difficult
Mission Operationsimplecomplexcomplexsimple

Design Selection

After careful consideration of all design alternatives, it was decided that design Option 4 with either a ramp or actuator platform should be developed further in the preliminary design stage. It is the most optimum design in terms of volume allowing for an increased payload size and ease of routing plumbing lines and wiring. The payload will need to be shielded since it is located above the main components and not buried within the spacecraft, but this is typically not difficult to do. This design is also quite attractive from the aspect of ease of conducting mission operations since both the oxygen plant and the rover are on the lunar surface. The primary challenge with this design will be in the design of the payload deployment system. As long as the mechanisms are kept as simple as possible and the CG remains approximately axisymmetric, this will not be a significant problem. This option meets all of the original design requirements and shows significant promise for success.

Design Development

Now that a conceptual spacecraft configuration has been selected, more energy can be devoted to developing a more detailed design layout. Several configurations have been developed using the Intergraph 3-D engineering modeling system. Three-dimensional solid modeling has proven to be an invaluable tool in the configuration analysis of the structural system. A three-dimensional model can show geometric and dimensional constraints which may cause problems for the design configuration. An initial idea of how the spacecraft might look is shown in Figures 5 through 8 below.

Figure 5 - Side View

Figure 6 - Bottom View

Figure 7 - Bottom View (Close Up)

Figure 8 - Spacecraft in Delta-II Payload Bay

Figures 5 and 6 show a nine-sided design that has the power and thermal control systems mounted in the base of the spacecraft between the propellant tanks. The GNC electronics and computers are located directly above the thermal control systems on a cold plate to maintain the operating temperature range (may not be necessary if a completely passive system is implemented). The rectangular volumes directly beneath the communications annena on the top of the spacecraft represent the rover and oxygen plant. This design divides the solar array area into two panels at 50W/m2 each and are located at an angle of 120 degrees. The side opposite the panels (the right side in Figure 6) is where the payloads will be deployed. Figure 7 shows a close-up of the base of the spacecraft illustrating one possible thrust structure concept and engine layout configuration. Figure 8 shows how the spacecraft fits into the volume requirements of the Delta-II payload bay with the landing gear and the solar panels retracted.

Future Plans

Now that the basic spacecraft and payload configuration has been determined, future efforts at Texas A&M will be focused on particular structural configuration of the spacecraft. This will include an analysis of the configurations described above as well as a determination of the number of sides on the spacecraft and the interface with the landing gear. The current thinking is that three legs will be employed and that the number of sides on the spacecraft will be some multiple of three to maintain axial symmetry and to facilitate a single type of interface with the landing gear. The methods of construction are also under consideration, with monocoque, skin-stringer, and truss designs all currently being evaluated. Work is also being done researching composite materials and how they can best be implemented for this design. Finite element analysis of each model will performed to determine the best load distribution and to improve the structural configuration.

The primary goals of the structures team in the project are listed as follows:


The initial effort in this project has been expended in determining an optimum spacecraft configuration that will be the best compromise of conflicting constraints and most efficiently meet the system design requirements. As the design develops and the systems are baselined, more detailed analysis will be implemented to more accurately determine the characteristics of the structure. This information will then be used to check that the design still meets all requirements and modifications will be made as necessary. This utilization of the iterative design process is critically dependent upon communication between the subsystem design teams and will maximize our chances at developing a successful design.


Wednesday, 31-Dec-1969 18:00:00 CST
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