SECTION IV - DEVELOPMENT - TEST - PRODUCTION

4.0 INTRODUCTION
Implementation of the Lunar Expedition will require a completely integrated program involving the development, test, and production of items based on almost every known technical discipline. These technical disciplines are presently being investigated under a multitude of programs and organisations. The Lunar Expedition program will require these technical efforts to be immediately organised and re-oriented where necessary. This can best be accomplished by preparing a detailed development, test, and production program. When this program is completed each technical area can be evaluated by comparing its present program objectives and its required output to meet the Lunar Expedition program requirements. In the following paragraphs the Lunar Expedition development objectives and technical performance requirements are presented. The scope of the major existing technical programs and the necessary re-orientation is discussed.

4.1 DEVELOPMENT OBJECTIVES
 

4.1.1 HIGH SPEED RE-ENTRY
At the present time high-speed re-entry data in the velocity spectrum from 25,000 ft/sec to 45,000 ft/sec is non-existent. In order to meet the Lunex Re-entry Vehicle development schedule it will be necessary to have high-speed re-entry data during the engineering design program for the manned re-entry vehicle. Thus a compressed and co-ordinated test program for both ground test facilities and flight testing is necessary.

Immediate action is necessary to schedule and design the high-speed wind-tunnel test program. This will show the type of information that can only be achieved by means of flight testing.

The High-Speed Re-entry flight test program scheduled for the Lunex program is necessary to provide basic data on re-entry as well as to fly specific shapes in the later period of the test program. This selected shape program will be co-ordinated with the Lunex Re-entry Vehicle design effort.

In order to accomplish the High-Speed Re-entry flight test program it will be necessary to design and develop a test vehicle. This vehicle must use existing boost systems due to time limitations, but the payload will have to be designed especially for this program since none exists at this time. It is believed that the Atlas booster will prove adequate for these tests, but a decision must await the test payload design.

4.1.2 MANNED LUNAR PAYLOAD
The largest single development objective for the Lunex program is to provide a payload capable of transporting men and equipment to the lunar surface and returning them to a selected earth base. This payload would consist of a Lunar Landing Stage, Lunar Launch Stage and a 3-man Lunex Re-entry Vehicle.

A typical Manned Lunar Payload is shown in a cut-away view in Figure 4-1. The characteristics and General Arrangement of the Manned Lunar Payload are seen in Figures 4-2 and 4-3. This payload is 52 feet 11 inches long, has the c.g. located 33 feet 8 inches from the nose of the re-entry vehicle and the interface diameter with the Space Launching System is 25 feet. The complete payload weighs 134,000 pounds at escape velocity, and a 20,205 pound Manned Re-entry Vehicle is returned to the earth.

The Lunex Re-entry Vehicle must be capable of entering the earth's atmosphere with a velocity of approximately 37,000 ft/sec. At the present time, basic re-entry information for velocities of this magnitude does not exist. Therefore, engineering design effort for this re-entry vehicle must be accomplished concurrently with other major sub-systems developments and integrated with the High-Speed Re-entry test program and the Abort System test and development program. This requires close management control of these programs by the Lunex Program Office.

Another major problem facing the re-entry vehicle development program is the life support package. The planned schedule will require the manned life support package to be designed on the basis of earlier primate shots, Mercury shots and the Discoverer series. These programs lead toward a manned capability, but this re-entry vehicle requires the first truly space life support package.

The Lunar Landing Stage must be capable of landing the Lunar Launching Stage and the Lunex Re-entry Vehicle on the lunar surface. At the present time this is considered a difficult design problem because little is known about the lunar surface. Actually the best photographic resolution to date is approximately 1/2 mile. Many theories exist on the formation of the moon and therefore, the characteristics of its present surface. When these two factors are considered the only practical design approach is to provide an alighting system capable of landing on an extremely rough surface. An automatic levelling, orientation and launching system is required for system check-out prior to manned flight. Therefore, any assumption that the Manned Lunar Payload can be moved about on the lunar surface or that the payloads might initially transfer fuel on the lunar surface, might be entirely erroneous and jeopardise the complete Lunar Expedition effort. The landing stage will also have to be developed so that it is capable of landing the Cargo Payloads on the lunar surface.

The Lunar Launching Stage must be developed with a different philosophy than the previous sub-systems. First, it only operates in the vacuum of space and on the lunar surface. Secondly, it will be required to function after it has been located on the lunar surface for an extended period varying from several days to many months. Therefore, the stage must be developed to launch the re-entry vehicle after being subjected to a better vacuum then available in our best earth laboratory facilities, following possible temperature variations of 400 to 500 degrees, following possible meteorite bombardment and from a less than optimum launch angle. Specifically the stage development must consider propellant boil-off, automatic check-out, self-erection and remote (earth-moon) launching procedures.

The Lunar Launching Stage represents the major reliability problem of the system because an abort capability is planned for every phase of the Lunex mission except during launch from the lunar surface. During the early lunar flights an abort capability for this phase is just too expensive payload-wise for the Space Launching System. An abort capability during lunar launch essentially requires a duplicate lunar launching capability because the man must still be returned to the earth by either this system, or a special rescue flight. Therefore, until lunar support facilities are available, a separate system for abort during lunar launch does not seem practical. This creates the requirement to develop an extremely reliable Lunar Launching Stage.

4.1.3 CARGO PAYLOAD
The successful support of the Lunar Expedition will require a capability to deliver relatively large Cargo Packages to the lunar surface. These Cargo Packages will be soft landed at the desired lunar sites by the Lunar Landing Stage. Each Cargo Package will weigh approximately 45,000 pounds and will be specifically designed to carry the items desired to support the expedition. Development of the Cargo Payload and the specific packages will depend upon the Lunar Landing Stage design and the receipt of lunar environmental data. The actual design of the Lunar Expedition Facility will only be possible when detailed information on the lunar surface is available. Then, with the facility design information, the required materials, equipment, and procedures can be determined and a payload delivery sequence derived. The required payload delivery sequence is essential before the individual payloads can be designed and developed, but timely development of major items of equipment must proceed as their individual requirements become known.

4.1.4 ABORT SYSTEM
The philosophy of abort has been presented in the Program Description section of this document. The development of the abort equipment will require an integrated effort with the re-entry vehicle design and the test program must be conducted concurrently to provide a reliable and safe system for supporting manned operations.

It is essential that the re-entry vehicle development be conducted so that the life support capsule can also meet the requirement imposed by the abort system. Additional structural and propulsion items must be developed to provide for abort during the earth ascent phase of the lunar mission. The computing and control equipment on the Manned Lunar Payload must be capable of selecting the desired abort mode of operation and initiating the desired actions at the required time throughout the lunar mission.

4.1.5 SPACE LAUNCHING SYSTEM
The Lunar Expedition requires an extensive space launching capability. The development of this capability is a necessary part of the Lunex Program. At present this development is being included under the Space Launching System program. It is designed to support the low altitude test, orbital, circumlunar, and full lunar flights.

One of the major problems facing the design and development of the Lunex payloads with reference to the Space Launching System concerns the interface characteristics, trajectory considerations, and earth launch facilities.

The present prime interface characteristics for the Manned and Cargo Lunar Payloads are as follows:


Interface Diameter - 300 inches
Escape Payload Weight - 134,000 pounds
Payload length - 635 inches
Center of Gravity (Measured from top of payload) - 404 inches

The Space Launching System is required to provide timely launching capabilities for the lunar Expedition as follows:

  • Payload Weight Pounds Trajectory Unmanned Flight Manned Flight

  • 20,000 300 mile orbit Aug 64 April 65

  • 87,000 300 mile orbit Dec 65 - -

  • 24,000 Escape Velocity Dec 65 Aug 66

  • 134,000 Escape Velocity July 66 Aug 67

4.2 SUBSYSTEM DEVELOPMENT

The development of a manned lunar payload and a cargo package requires the development of subsystems and applied research in many technical areas. Studies have established that the advances in performance in these technical areas can be accomplished to meet the overall program schedulers and that no scientific breakthroughs are required. The important point is that items requiring development be identified, that necessary funds be allocated, and that effort be initiated without delay. The following sections discuss major subsystem requirements, present capabilities, and development required. Completed studies conducted by the Air Force and industry have established subsystem requirements in sufficient detail to outline development programs which should be initiated immediately. Present studies will refine these specifications further.

4.2.1 RE-ENTRY VEHICLE

4.2.1.1
The manned re-entry vehicle is a critical item in the development of the manned payload packages. This vehicle must be capable of returning from the moon and re-entering the earth's atmosphere at earth escape velocity (37,000 ft/sec). It must also have the capability of supporting three men on a 10-day round trip earth-moon mission. This mission would include boost from earth, coasting in earth orbit, ballistic flight to the moon, de-boost and landing on the moon's surface, remaining on the moon for one to five days, launch from the moon's surface, re-entering the earth's atmosphere and landing at a pre-selected base on the earth.

 

Structural requirements imposed by inertial and pressure loading during boost, abort, trajectory correction, landing, re-entry, ground handling, and wind loading on the launch pad, have been considered in analysing desired vehicle characteristics. These studies have also included the heating and its effect on vehicle design as well as the effects of space and lunar environment including particles and radiation, meteorite penetration, and hard vacuum. Present design studies have estimated the total re-entry vehicle height at 20,205 pounds. The weight breakdown is as followers:

a. Body 7500

(1) Structure 3500
(2) Heat Shield 4000

b. Wing Group 2000

(1) Structure 800
(2) Heat Shield 1200

c . Control System 775

(1) Aerodynamic 600
(2) Attitude 175

d. Environmental Control 1530

(1) Equipment Cooling 138
(2) Structure Cooling 940
(3 ) Cryogenic Storage 452

e. Landing Gear 700
f. Instruments & Displays 200
g. Electric Power System 600
h. Guidance & Navigation 400
i. Communications 250
j. Furnishings & Equipment 850

(1) Seats & Restraints 225
(2) Decompression Chamber 175
(3) Equipment Compartment 300
(4 ) Miscellaneous 150

k. Life Support 400
l. Crew (3 men) 600
m. Radiation 1200
n. Abort System 3000

4.2.1.2
Present re-entry and recovery techniques are outgrowths of the ballistic missile program utilising ballistic re-entry and parachute recovery. They are not compatible with the velocities associated with re-entry from the moon, with controlled landing, or with manned operations. Present engineering data associated with high speed re-entry is not adequate for vehicle design.

4.2.1.3
A development test program is required to obtain generalised data on re-entry phenomena and to test scale models of selected vehicle configurations so that final selection and design of an optimum vehicle can be made. Concurrently with this test program the projects within the applied research program will be directed so as to carry out the following investigations to provide necessary data for the Lunex Re-entry Vehicle Design.

4.2.1.3.1 AERODYNAMICS

(1) Study hypersonic low density aerodynamics including dissociation and ionisation, non-equilibrium flow phenomena, and the influence of radiation non-equilibrium on vehicle aerodynamic and heat transfer characteristics.

(2) Initiate an extensive ground based facility program directed at obtaining aerodynamic and heat transfer data up to Mach No. 25 (the maximum useable available capability). These tests would include the G.E. hypersonic shock tunnel in the M = 18 - 25 range; Cornell Aeronautical Laboratory hypersonic shock tunnel M = 12 - 18; Cornell Aeronautical Laboratory heated hydrogen hypersonic shock tunnel at M = 20; AEDC tunnel "B", "C", at M = 8 - 10; AEDC E-1 and E-2, M = 1.5 - 6; AEDC supersonic and subsonic facilities. This effort will be co-ordinated with the Lunex Engineering Design program and the High-Speed Re-entry test program.

(3) Correlation of wind tunnel tests in terms of prediction of free-flight vehicle performance characteristics in order to provide correlation between ground test facilities and free-flight vehicles.

(4) Complete vehicle static and dynamic stability analysis.

(5) Investigate local critical heat transfer problems including those associated with flaps and fins. The use of reaction controls, in order to alleviate critical heating areas, for vehicle stability and control, will be investigated.

4.2.1.3.2 MATERIALS

(1) Materials Development

(a) Low conductivity plastic material development
(b) Uniformly distributed low conductivity.

(2) Tailoring conductivity distribution in material in order to obtain high ablation performance at surface and low thermal conductivity in structure bond line.

(3) Develop materials with low ablative temperatures.

(4) Investigate bonding of materials to hot structure.

(5) Develop minimum shape change materials for aerodynamic control surface and leading edge applications. These materials will include pyrolytic graphite, alloys of pyrolytic graphite, and ceramics.

(6) Materials Analysis

(a) For selected materials above, develop analytical model to predict ablation performance and insulation thickness.
(b) Experimentally study material performance under simulated flight environments with the use of high enthalpy arc facilities (h/RT-0 = 700 to 800).
(c) Study the influence of space environment on selected materials. This will include the influence of vacuum, ultraviolet radiation, and high energy particles.

4.2.1.3.3 STRUCTURES

(1) Primary effort will be in the development of load-bearing radiating structures. For this structure, the following areas will be investigated.

(a) Thermal stress analysis and prediction.
(b) Dynamic buckling
(c) Strain gage applications to high temperatures.
(d) Experimental simulation on large scale structures of load temperature distribution, and history. The WADD Structures facility would be the one most appropriate to these tests.

4.1.2.3.4 DYNAMICS

(1) Analytical studies in the following areas should be undertaken.

(a) Unsteady aerodynamic forces at hypersonic speeds.
(b) Aeroelastic changes in structural loading and aerodynamic stability derivatives.
(c) Flutter
(d) Servoelastic coupling with guidance system.
(e) Fatigue due to random loading.
(f) Transient dynamic loading.

4.2.1.4
Present projects within the Air Force applied research program will be reviewed and reoriented or effort increased, as appropriate, to provide the necessary data. Projects which can be used for this purpose are listed below:

  • 6173 (U) Study of Controlled Final Deceleration Stages for Recoverable Vehicles.

  • 1315 (U) Bearings and Mechanical Control Systems for Flight Vehicles.

  • 1368 (U) Construction Techniques and Applications of New Materials.

  • 1370 (U) Dynamic Problems in Flight Vehicles.

  • 1395 (U) Flight Vehicle Design.

  • 6146 (U) Flight Vehicle Environmental Control.

  • 1309 (U) Flight Vehicle Environmental Investigation.

  • 6065 (U) Performance and Designed Deployable Aerodynamic Decelerations

4.2.1.5

In addition to the applied research efforts referred to in Paragraph 4.2.1.4 an intensive study of re-entry vehicle characteristics required for the Lunex mission is being accomplished under project 7990 task 17532. This study will define an optimum vehicle configuration and present the most feasible technical approaches to solving the various re-entry problems. For example, the desirability of ablative and/or radiation techniques for cooling will be determined.

4.2.2 PROPULSION

4.2.2.1
The Manned Lunar Payload requires a booster capable of placing a 134,000 pound package at escape velocity on a selected lunar trajectory. This booster development has been included in the Space Launching System Package Plan and its development will be done for the Lunex program.

4.2.2.2
Propulsion systems for the Manned Lunar Payload which will be developed under this plan are those required for the following operations:

  • Lunar Landing

  • Lunar Launch

  • Trajectory correction

  • Attitude control

  • Abort

4.2.2.3
The Lunar Lending Stage must be capable of soft landing at approximately 20 ft/sec a 50,000 pound payload on the moon. This payload consists of the Lunar Launching Stage and Lunex Re-entry Vehicle. Preliminary design data from studies completed to date show that the manned re-entry vehicle will weigh approximately 20,000 pounds and a launch stage of 30,000 pounds will be required. Similar estimates for the Lunar Landing Stage indicate that it will weigh 85,000 pounds. During lunar landing, if an initial thrust to weight ratio of 0.45 is assumed as consistent with the deceleration desired and time of deboost, an initial retro thrust of 60,000 pounds is required. At final touchdown on the moon, with all delta-v cancelled and assuming essentially all de-boost propellant consumed, approximately 10,000 pounds of thrust is required. Some throttling or gimballing of the engine may be required at the 10,000 pound level to reduce the axial component of thrust. The requirements on the landing engine are for a 60,000 pound engine with a 6 to 1 throttling ratio, or a cluster of four engines of 15,000 pounds thrust and at least one with a throttling range of 1.5 to 1. Assuming a thrust to weight ratio of 1.5 (Moon weight) for the Lunar Launch Stage, a 12,000 pound thrust engine is required for lunar launch. An engine of the LR-115 type will meet these requirements with some development. Minor development will be necessary if the range of throttleability is 20 to 30%. If the range of thrust control is 50% or greater, a more extensive program will be required.

4.2.2.4
In addition to the deboost and launch, it is necessary to provide a trajectory and attitude control propulsion capability. A velocity capability of 300 to 1200 ft/sec will be required for trajectory corrections during midcourse, lunar landing and return. Attitude control will be required during lunar landing and launch, and midcourse, with specific methods to be determined by optimisation studies during vehicle design. There do not appear to be any major development problems to be overcome to provide trajectory correction or attitude control capability.

4.2.2.5
An abort system to provide safe removal of the crew in the event of failure before, or during launch must be developed. A propulsion system with an extremely short reaction time is necessary to insure safe crew removal.

4.2.2.6
Specific engine sizing, throttleability requirements, propellant and oxidizer selection, nozzle type, etc., will be determined upon completion of a preliminary design in which such tradeoff comparisons as range of throttling versus use of verniers will be made and optimized selections made. Development work will be initiated within present projects in the Air Force applied research program to raise the level of technology in areas such as throttleability. Projects which can be utilized for this purpose are:

  • 3085 (U) Liquid Rocket Engine Technology

  • 3148 (U) Development of Liquid and Solid Rocket Propellants

  • 6753 (U) Rocket Propulsion Subsystems

  • 6950 (U) Propulsion Attitude Testing

4.2.3 LIFE SUPPORT

4.2.3.1
The life support package for the manned Lunar Payload will be required to function for a minimum of 10 days. This is based on the premise that a one-way trip to the moon will require 2 1/2 days, and the stay on the lunar surface will be on the order of 5 days. The life support systems must be capable of supporting three men during high acceleration boost, approximately 2 1/2 days of weightlessness, one to five days of 1/6 earth weight, 2 1/2 days of weightlessness, re-entry deceleration and return to full earth gravity. At the same time it must provide a shirtsleeve cabin environment under the space and lunar environments, including extreme temperature gradients, absence of oxygen, radiation, etc.

4.2.3.2
Studies of the life support system weight requirements indicate that the life support package can be provided within the weight allocation for the 20,000 pound Lunex Re-entry Vehicle. The life support system weight analysis was based on physiological experiments under simulated apace flight conditions such as confinement, special diets, reduced pressure, etc. At the present time approximately 65 to 70 percent of the knowledge required to design the three man package is available. However, to obtain the additional data experimental laboratory and flight testing is required. Most information is presently obtained by piggyback testing aboard experimental vehicles, but to support the Lunex program and to meet the desired schedules the BOSS primate program must be initiated and adequately supported.

4.2.3.3
Most of the data available today consists of physiological support (nutrition, breathing oxygen, pressure suits, and restraints for limited periods), but there is a lack of knowledge on prolonged weightlessness and the biological effects of exposure to prolonged space radiation. The BOSS program initially will support a chimpanzee for periods up to 15 days and has been programmed to provide a life support package of sufficient size and sophistication to support a man. Thus, with the BOSS program the data will become available so that the Lunex program can design and construct the life support package as required for April 1965.

4.2.3.4
Throughout this development all life support knowledge and techniques will be fully exploited. Techniques learned in the work with the Discoverer package were utilised in building the Mercury package. In turn, experience and knowledge gained from Mercury is being fully exploited in the development of the present BOSS package .

4.2.3.5
The life support program (BOSS) is vital to meet the objectives of the Lunex program. However, other AFSC programs must be considered for possible application to Lunex and the following are now being evaluated:

  • 6373 (V) Aerospace Life Support

  • 7930 (U) Bio-Astronautics

4.2.4 FLIGHT VEHICLE POWER

4.2.4.1
Electrical power will be required to operate the Lunex Re-entry Vehicle subsystems such as life support, navigation and guidance, instrumentation, and communications. The power requirement for there subsystems, has been analysed and determined to he approximately 3 kW average during a ten-day manned trip to the moon and return. Peak power requirement will be approximately 6 kW.

4.2.4.2
Solar, nuclear, and chemical powered systems were evaluated against these requirements. While all of these systems may be capable of meeting these requirements the chemically powered systems have been selected for early adaptation into the program. Specifically, fuel cells and turbines, or positive displacement engines appear to offer the moat advantageous solution. The final selection will be made during the final re-entry vehicle design when a detailed analysis of the trade-off between various available systems considering relative weight, efficiency, reliability, and growth potential is available. The optimum system may be a combination of fuel cells and chemical dynamic systems with one system specifically designed to supply peak demand. With this approach the system to provide peak load capacity, will also provide backup power in the case of equipment malfunction during a large part of the mission. A battery supply may be used to furnish emergency power required for crew safety during critical periods in the flight.

4.2.4.3
Present level of technology is such that a satisfactory flight vehicle power system will be available when required for the Lunex mission. Additional development effort should be initiated in certain specific areas, such as a reliability evaluation program for fuel cells and an investigation of the problems of operating chemical dynamic systems in the zero G environment.

Close co-ordination must also be maintained with the manager of project 3145 (U) Energy Conversion, to insure the availability of the required secondary power sources.

4.2.5 GUIDANCE AND CONTROL SYSTEM

4.2.5.1
A study of the guidance and control requirement for the lunar vehicle indicates that the mission can be accomplished by reasonable extensions of present state-of-the-art equipment. The complete lunar vehicle guidance package should be capable of furnishing guidance and control during the following phases of the lunar mission.

  • Ascent and Injection

  • Outbound Mid-course

  • Lunar Landing

  • Lunar Ascent

  • Inbound Mid-course

  • Earth Re-entry

  • Earth Landing

Present state-of-the-art equipment is capable of handling portions of the guidance and control problem associated with the above phases of flight. However, in order to obtain a complete guidance and control system, it is felt that development of the following items should be undertaken.

4.2.5.2 INERTIAL PLATFORM
Guidance requirements for both the manned and unmanned vehicles can be met with the use of guidance concepts based on the use of inertial and corrected inertial data in a combination of explicit and perturbation computations of present and predicted trajectories. Consequently, an inertial platform configuration suited to the space environment is needed. This platform should be light in weight, highly reliable, and capable of maintaining a space-fixed reference over a long interval of time. Present gyroscopic devices and accelerometers are neither accurate nor reliable enough to accomplish the space mission.

An inertial platform which holds great promise for use in lunar missions is one utilising electrically suspended gyros in conjunction with advanced accelerometers capable of operating in a space environment. Present electrically suspended gyros are capable of operating with a drift rate of .0005 deg/hr/g and it is anticipated that by 1966, a drift rate of .0001 deg/hr/g will be attainable. Also, no difficulties are foreseen in maintaining suspension of the rotating member in an acceleration field of 15 G's with 30 g's being possible. Development of a small inertial platform utilising electrically suspended gyros will be required for the lunar mission.

4.2.5.3 STAR TRACKER
In order to increase the reliability and the accuracy of the inertial platform, a compact star tracker for use with the platform during the outbound and inbound mid-course phases of the lunar flight is desired. Also, the star tracker should be capable of operating in a lunar environment so that it can be used for stellar alignment during the lunar launch portion of the mission. The accuracy of present solid state star trackers is approximately 10 seconds of arc and their weight is approximately 15 pounds.

 

However, these trackers are untested in a space environment and must be developed for the lunar mission and for use with the small inertial platform. In particular, the star tracker must be capable of furnishing accurate stellar alignment information to the inertial platform during the lunar ascent portion of the mission. If it is possible to develop a controllable thrust engine in time to meet the launch schedule, the boost and injection guidance problem for the lunar ascent will be simplified as it will be possible to time-control a predetermined velocity path. This development could possibly reduce the accuracy requirement of the star tracker.

4.2.5.4 LONG BASELINE RADIO NAVIGATION
Since manned as well as unmanned flights are planned for the lunar mission, it is necessary to have a navigation system to back-up the inertial system and to increase the over-all accuracy of the guidance and control techniques. Long baseline radio/radar tracking and guidance techniques offer great possibilities for tracking and guiding vehicles in cislunar apace.

 

Present studies show that there are a number of problems yet to be solved to give the long baseline radio navigation the desired accuracy. Among these problems are 1) the accuracy with which co-ordinates can be determined for each tracking station, 2) the accuracy with which corrections can be made for tropospheric and ionospheric propagation effects on the system measurements, and 3) the accuracy with which "clocks" can be synchronised at the several stations. Reasonable extensions of the state-of-the-art should be able to overcame these problems however, and it is felt that the development of a long baseline radio navigation system will be necessary for the lunar mission.

4.2.5.5 DOPPLER RADAR
Anticipation that radio beacons will be in place on the lunar surface has somewhat simplified the lunar landing phase of the mission. The use of mid-course guidance will enable the vehicle to approach the moon within line-of-sight of at least one of the radio beacons, and the beacon can be utilised for the approach phase of the lunar landing. However, for final vertical velocity measurement, a sensing technique particularly sensitive to small velocity changes is required. A small CW Doppler radar is ideally suited for this requirement. Therefore development of a small, reliable Doppler radar which can operate in the lunar environment is needed. In order to decrease the power requirement for the radar it should not be required to operate at a range of over 300 miles.

4.2.5.6 RE-ENTRY GUIDANCE
Major emphasis must be placed on the guidance requirements for the re-entry phase of the lunar mission. Position, velocity, and attitude can be measured by the inertial system, however, other measurements initially required will be temperature, temperature rate, structural loading and air density. Extensive further study is needed to define these measurements with any accuracy. Early earth return equipment should furnish the data necessary to develop the required re-entry guidance package for the lunar mission.

4.2.5.7 ADAPTIVE AUTOPILOT
The control of the re-entry vehicle mill require an adaptive autopilot due to the wide variation in surface effectiveness. Adaptive autopilots such as used in the X-15 are available, but extensive development is needed to ready them for use in the lunar mission.

4.2.5.8
The following projects or specific tasks within these projects can be utilised to provide the development required for the Lunex program.

  • 4144 (U) Guidance and Sensing Techniques for Advanced Vehicles

  • 40165 (U) Data Conversion Techniques

  • 50845 (U) Guidance Utilising Stable Timing Oscillators

  • 50899 (U) Molecular Amplification Techniques

  • 4427 (U) Self-Contained Electromagnetic Techniques for Space Navigation

  • 4431 (U) Inertial System Components

  • 44169-II (U) Space Adapted Celestial Tracking System

  • 44169-III (U) Multi-Headed Solid State Celestial Tracker

  • 44169-IV (U) Solid State Celestial Body Sensors

  • 5201 (U) Inertial Systems Technique

  • 5215 (U) Military Lunar Vehicle Guidance

  • 50820 (U) Military Lunar Vehicle Guidance Systems

  • 58821 (U) Military Lunar Vehicle Terminal Guidance

4.2.6 COMPUTER

4.2.6.1
The United States has the ability to provide a suitable computer facility at the present time to support the Lunex mission. As the milestones in the program are realised and requirements become more complex, the computer capability will improve to meet these more stringent requirements. Detailed studies on the specific needs of the missions, time-phased, will be conducted to determine trade-offs among possible techniques to insure that machine sophistication does not became an end unto itself. The following guidelines providing adequate flexibility, have been followed in arriving at the required development recommendations:

a. Manned vehicles will require extensive data reduction to give an operator real-time displays of the conditions around him and solutions to problems such as, velocity and attitude corrections, etc.
b. Sensor control (aiming and sampling rate) and data processing will be accomplished on the vehicle either on ground command, or by operator direction.
c. Mid-course and terminal guidance requirements will make severe demands upon vehicle-borne computational systems.
d. Radiation hazards and effects which are unknown at present could influence the technology that will be utilised for lunar missions.
e. Emergency procedures must be available in the event that the operators became incapacitated and incapable of returning to earth at any time during the mission.

4.2.6.2
The computer capability can be expanded in two basic ways by improved hardware, or new concepts. Examples of new approaches which will be reviewed prior to selection of the final vehicle design are the following:

a. Standardised computer functions incorporated into modules so that they can be used to "build" their capability for each mission required. Such a concept would allow a vehicle designer to fabricate a computational facility without resorting to extensive redesign and/or re-packaging. The modularised concept noted above is particularly adapted to unmanned missions.

b. For a manned mission two fixed programs could be permanently placed in storage; these would be an overall command, or executive routine to direct the sequences of operation, and the other could be an emergency return-to-earth routine that could be actuated by the master control. Thus a 5-pound tape unit would replace a larger core memory and provide a higher degree of flexibility. The principal advantage in this system is that the computer is general-purpose in design and therefore useable on a large variety of missions and unnecessary capabilities will not be carried on a particular mission.

c. An optimised hybrid of analogue and digital devices combined to use the better features of each, i.e., speed of problem solution from the analogue and precision, flexibility, and data reduction from the digital.

4.2.6.3
Substantial improvements in computer capability, developments, reliability, volume, weight, and power consumption will be available for the Lunex program by effort expended in the following areas:

a. Core-rope memories to be used in fixed memory applications.

b. Functional molecular blocks. By 1963, the date of earth orbit, it is expected that more than 80% of all computer functions can be performed by this method. Advantages are numerous: high memory densities, extremely small size, small weight and power consumption.

c. Self-healing, or adaptive programming techniques as a means for back-up on component reliability.

d. Electroluminescent-photoconductive memory devices should be considered for their radiation and magnetic invulnerability. In this regard, pneumatic bistable elements should be considered for the same reason.

e. Photochromic storage devices have advantages in high storage densities, 1 billion bits/cubic inch. Certain applications, such as semi-permanent storage, could benefit from this feature.

4.2.6.4
The following projects in the Applied Research Area will be utilised to obtain improvement in computer technology:

  • 3176 (U) Space Borne Computation & Control Techniques

  • 4421 (U) Digital Computation Methods & Techniques



DEVELOPMENT - TEST - PRODUCTION
 

4.2.7 COMMUNICATIONS

4.2.7.1
The manned lunar mission will require communications channels between the vehicles and earth and on the lunar surface for telemetry, T.V., voice, and vehicle control. Specific system parameters will depend on the characteristics of the ground tracking network and communications stations which will be used to support the lunar missions.

4.2.7.2
There are no significant technical problems associated with the development of equipment to perform the required communications operations. One exception to this general statement is that during re-entry radio transmission may not be possible at the lower frequencies utilised elsewhere in the mission because of the plasma shield set up by aerodynamic heating. One possible solution may be to provide a separate system operating at 10,000 mcs for re-entry. Overall savings in equipment weight, and power requirements will result from careful analysts and identification of requirements for information transfer and maximum utilisation of system components in a dual role. This will be done during the vehicle design phase. While not a requirement for early missions the capability to provide a secure communications link is desirable and will be considered during final design of the communications systems. A secure communications link will be a requirement in later missions. Throughout all phases, communications links critical to mission success should incorporate a high degree of protection against natural or man-made interference, or deliberate jamming.

4.2.7.3
The following Air Force projects will be reviewed and used to provide the necessary results required for the Lunex mission:

  • 4335 (U) Applied Communications Research for Air Force Vehicles

  • 4519 (U) Surface & Long Range Communications Techniques

  • 5570 (U) Communications Security Applied Research

4.2.8 ENVIRONMENTAL DATA

4.2.8.1
Present knowledge of the lunar environment is extremely limited and it is necessary to obtain detailed information concerning the lunar composition, subsurface structure, surface characteristics, meteorite flux, level of solar and cosmic radiation, and magnetic field. This knowledge is required to design the equipment for the Lunex program so that personnel may be protected and the mission accomplished.

4.2.8.2
The importance of lunar composition in manned exploration of the moon lies largely in the ability of the moon to provide fuel for vehicles and secondary power, as well as to supplement life support systems with additional water, radiation shielding material, and semi-permanent shelters. Of these lunar resources, water appears to be of major importance both as a fuel and in life support. Water will probably be present both as ice in permanently shadowed zones and as water of hydration in certain minerals such as serpentine.

4.2.8.3
Present knowledge of lunar composition is almost entirely theoretical. The relatively low lunar density (3.34) indicates low metal content. By analogy with the compositions of meteorites it is generally assumed that the moon is composed of chondritic (stony meteorite) material. That this assumption is only partially valid is demonstrated by the fact that chondritic meteorites would have to lose about 10% of their iron content in order to attain this lunar density.

4.2.8.4
The Air Force and NASA are presently trying to determine the lunar composition indirectly through study of tektites, which may be fragments of the moon, and through study of micrometeoritic dust captured above the atmosphere. (Air Force efforts are funded under Project 7698).

4.2.8.5
The Air Force is trying to determine the lunar composition directly by means of spectrometric analysis of the natural X-ray fluorescence of the moon due to the bombardment of the lunar surface by solar radiation. The first knowledge of lunar composition is anticipated in March of 1962. (This work is also funded under Project 7698).

4.2.8.6
NASA intends to measure the lunar composition directly by means of its Surveyor lunar probe now scheduled for mid-1963.

4.2.8.7
Neither Air Force measurements of overall lunar composition, nor NASA measurements of spot compositions will satisfy the requirement for location of lunar resources. The NASA Prospector vehicle scheduled for 1966 will obtain more widespread data, but that is urgently needed is detailed knowledge of the variation of lunar composition over the whole surface. This can only be accomplished by a lunar orbiting vehicle with appropriate instrumentation. NASA presently has this planned for 1965 and the appropriateness of their instrumentation remains in doubt. Also this is too late to meet the requirements of the Lunex program.

4.2.8.8
The importance of lunar subsurface structure in exploration of the moon lies largely in a possible collapse hazard under vehicles and personnel, and in the possibility of utilising subsurface structures as shelters and storage facilities.

Present knowledge of lunar subsurface structure is based on a theoretical extrapolation from the presumed origin of the surface features. The majority of lunar geologists believe that lunar craters were formed by means of the impact of large meteorites, and that only limited volcanism has occurred in the lunar highlands. The maria, on the other hand, are thought to be giant lava pools; although the melting is assumed to have been triggered by asteroidal impact.

Based on these theories of origin for the lunar surface features, it is thought that the subsurface structure of the lunar highlands will consist largely of overlapping layers of debris ejected from the impact craters. The collapse hazard of such material is negligible. The maria should be covered by no more than 40 feet of vesicular (bubble filled) lava, with maximum vesicle (bubble) size about six feet in diameter. Such terrain could present a collapse hazard, the severity of which will depend upon actual (rather than maximum) vesicle size.

It should be noted, however, that a rival theory for the origin of lunar craters holds that they were produced by volcanism as calderea. Should this theory be correct, the collapse hazard in the highlands would probably exceed that on the maria.

In order to determine the lunar subsurface structure, it is necessary to place instruments on the moon. Thus, the Air Force, although contributing theoretical evaluations an described above (under Project 7698), has no program for directly determining lunar subsurface structure. NASA plans to place seismometers and a coring instrument in the Surveyor vehicle in mid-1963 to determine these parameters. Again, point measurements are not sufficient, and geophysical instrumentation adequate for determining subsurface structure from the lunar orbiting vehicle (1965) should be developed.

4.2.8.9
The importance of lunar surface characteristics lies in their critical importance in design of both rocket end surface vehicles and in lunar navigation. Critical surface characteristics include gross topography, microtopography, and the nature of the lunar dust. Of these characteristics, knowledge of gross topography will be important in overall rocket design and in design and operation of rocket landing and navigational equipment. The microtopography (relief less than 20 feet) will be important in the design of rocket landing equipment and the vehicle for surface exploration. The nature of the surface dust will be moat important in design of the vehicle for surface exploration.

Present knowledge of gross topography shows that slopes are generally gentle, and topographic profiles have been determined over a limited amount of terrain. Present knowledge of microtopography is very limited. Radar returns, once thought reliable indicators of low microrelief, are now considered by moat space scientists to be so poorly understood that conclusions may not be drawn from them. Photometric data appears to indicate a rather rough surface, but this data is also subject to more than one interpretation. Present knowledge of the nature of the lunar dust is entirely theoretical. The leading school of thought holds that the dust is compacted and sintered. An opposing school holds that the dust bears an electrostatic charge. Should the dust bear an electrostatic charge, it would be very loose and probably subject to migration. The hazard to surface vehicles and even personnel is apparent.

Gross lunar topography on the visible face is presently being mapped by the Aeronautical Chart and Information Center based on techniques developed under Project 8602. Maximum resolution is about 1/3 mile, and average resolution is about one mile. Higher resolution photography and photography of the back side of the moon will be obtained by the lunar orbiting vehicle planned by NASA for 1965. A co-operative effort by ACIC and NASA is presently envisioned to produce the necessary topographic lunar charts.

Microtopography is being studied by the Army Corps of Engineers through radar experiments. (the Air Force work is being done on the Millstone radar equipment) The nature of the lunar dust is being studied primarily by the Air Force under Projects 7698 and 8602 by means of radiometrlc studies from high altitude unmanned balloons and results are anticipated in early 1962. NASA anticipates obtaining at least partial data on the nature of the dust from Surveyor (mid-1963) by television observation of the lunar surface and by the landing characteristics of the vehicle.

4.2.8.10
The meteorite flux and level of solar and cosmic radiation near the lunar surface are important for the survival of personnel either on the lunar surface or in vehicles and shelters.

Present knowledge of these parameters is fairly precise as a result of satellite and deep space probe experiments by NASA and the Air Force. Only the radiation environment within the first few meters of the lunar surface is still speculative as a result of uncertainties in our knowledge of the interaction of solar and cosmic radiation with the lunar surface materials. It seems likely that a cloud of ions will be produced by radiation bombardment as well as secondary X-rays. The density of the electron cloud is unknown, and may be critical for lunar communications.

The Air Force is studying the lunar and cislunar radiation environment under Projects 6687, 6688, 7601, 7649, and 7663 by means of satellites, deep space probes, and vertical sounding rockets. The NASA Surveyor vehicle (mid-1963) should give detailed knowledge of the radiation environment at the lunar surface.

4.2.8.11
The lunar magnetic field may be important to space and lunar surface navigation, and in its effects on ionised lunar materials.

The Russian Lunik II indicated that the lunar magnetic field must be very small. The Russians were not clear on how small, but it is generally thought that the moon does not possess a magnetic field. Thus, all magnetic effects should be derived from the very low intensity interplanetary field and magnetic fields, "frozen" into solar plasmas.

The Air Force is studying the interplanetary magnetic field under Project 7601. NASA should be able to determine the field near the lunar surface by means of the Ranger vehicle during 1962, and the field at the surface by means of Surveyor (mid-1963).

4.2.8.12
Two facts are apparent from a study of the data outlined above. First, one Air Force project (7698), which was funded for only 250K in 1962 and 300K in 1963, is responsible for moat of the research on the lunar environment. More funds are required for a speed-up in this program. Second, many critical experiments are tied to the NASA lunar orbiting vehicle, which has not been considered a very important vehicle by NASA. This program is essential to the Lunex and should be speeded up and planned more carefully.

4.2.8.13
Effort on Air Force Project 7698 will be increased substantially to insure that data necessary for the Lunex mission is available. A close working relationship will be established with NASA to provide, if possible, the inclusion of Air Force requirements in their planning for lunar programs both as regards to objectives and scheduling.
 


4.2.9 MATERIALS

4.2.9.1

The lunar expedition imposes rigid requirements on materials to maintain their characteristics while subject to radiation, vacuum, temperature extremes, and meteorites. This problem must be considered by the individual subsystem design. It is intended to point out here the overall material problem and programs which will contribute to its solution.

4.2.9.2

The absence of an atmosphere on the moon increases the radiative flux (particle and electromagnetic) from the sun and as such potentially increases the possibility of damage to man and lightweight plastic structures through the formation of free radicals and subsequent depolymerisation. The need for lightweight shielding is apparent. The vacuum conditions of the moon would aggravate the problems associated with moderately volatile constituents of plastics, lubricants, etc. For instance, the relatively volatile plasticizers in a plastic material could evaporate and interfere with the plastic function. Finally, the results of impact of micrometeorites on structural materials must be determined. All desirable properties must be acquired without penalty of weight. In addition to the problems encountered on the Moon, similar problems are encountered while in transit. In particular the heating encountered on re-entry into the Earth 'e atmosphere at 37,000 feet per second presents a severe material problem.

4.2.9.3

Some of the specific material requirements that can be identified are:

a. Lubricants that will function for long periods of time in a vacuum and temperature conditions such as exist in the moon.
b. Materials that will not sublimate in a vacuum at moon temperature.
c. Light-weight shielding material against meteorites.
d. Light-weight radiation shielding.
e. Shock-absorbing material that will function at 330 degrees F.
f. Coatings that will resist radiation, especially during periods of solar flares.
g. Glues and adhesives that will function with lunar materials.

4.2.9.4

Present projects to raise the level of technology in materials are listed below. They will be supported as required to insure success of the lunar mission.

  • 7312 (U) Finishes and Materials Preservation.

  • 7320 (U) Air Force Textile Materials.

  • 7340 (U) Non-Metallic & Composite Materials.

  • 7351 (U) Metallic Materials.

  • 7371 (U) Applied Research in Electrical, Electronic, and Magnetic Material.

  • 7391 (U) Energy Transmission Fluids.

4.2.9.5

While work in the basic research program cannot be counted on to provide technical breakthroughs within the time schedule of the Lunex program, materials study of this type will be monitored so that all technical advances can be integrated into the Lunex program. Specific examples of projects of this type are:

  • 8806 (U) Research on Materials at High Temperature.

  • 7022 (U) Surface and Interface Phenomena of Matter.

  • 9760 (U) Research in Properties of Matter.

 

4.3 TEST PLAN


The development and production of the equipment for the Lunar Expedition will require a concurrent and detailed test program.

The test program will be carried out on the basis of research tests to establish design criteria, materials tests, component tests, and finally, a progressive series of tests as components are assembled into subsystems and major systems and structures. Integration tests for flight suitability will be conducted for all functioning systems and the complete vehicle. Payload effects on the booster structure will be determined with a simulated payload. Subsequently, a flight-type payload will be used to demonstrate booster-payload system compatibility, reliability, crew safety, and mission performance.

Emphasis will be placed early in the program on research tests to derive basic design criteria, define the configuration and determine aerodynamic parameters.

Tests are to be run at progressively higher levels as the design evolves. Thus, entire subsystems, combined subsystems and complex major structures are to be subjected to evaluation tests as necessary to investigate component and subsystem interactions, or to prove out complex structural designs.

A captive test vehicle firing program will be the culmination of ground development testing. The over-all objective of the captive-firing program is to demonstrate satisfactory integration of the propulsion system with other vehicle systems that have an interface, direct or indirect, with the propulsion system. The early tests will be conducted in a simulated vehicle with the airborne vehicle systems installed on a heavy-wall propellant tank section. The tanks will be supported by a test stand structure which will also restrain the tanks against propulsion system thrust forces. For final testing a flight-type configuration will be used during captive tests.

Flight testing of the High-Speed Re-entry Test Vehicle, the Abort System, and Orbital, Circumlunar and unmanned lunar landing and Return Vehicles will complete the development program.

4.3.1 TEST CATEGORIES

4.3.1.1 RESEARCH TESTS
Tests will be run in appropriate research laboratories to define basic design criteria in at least the following technical areas:

a. Propulsion
b. Heat transfer
c. Aerodynamic forces and pressures
d. Materials
e. Statics (structures)

4.3.1.1.1 Propulsion Tests

Wind and vacuum tunnel testing will be conducted to investigate the problems of multiple re-start in a vacuum environment, to develop throttleable techniques, to determine lunar landing problems, and to determine the desirability of using the same engines for lunar landing and lunar launching.

Tests will be made to evaluate the propulsion stage for the circumlunar flights and to determine the capability of the abort propulsion system to accomplish its objective.

4.3.1.1.2 Heat Transfer Tests

Testing will be required on the insulation for the liquid hydrogen tanks to determine:

a. Optimum material thickness and weight
b. The amount of liquid hydrogen boil-off
c. the air leakage through seals
d. The airload effect on structural integrity
e. The thermal bowing of insulation panels
f. The separation distance between panel and tank skin

Scale-model or modified full-scale air-conditioning tests will be conducted on engine components, adapter sections and flight equipment storage areas.

Heat transfer characteristics for selected materials, structures, and surfaces will be required to support the engineering design.

4.3.1.1.3 Aerodynamic Force and Pressure Wind-Tunnel Tests

Wind-tunnel model tests of the launch vehicle and payload configuration will be required to accurately determine the aerodynamic forces and moments imposed on the vehicle during the boost trajectory.

These tests will provide data for determination of structural design criteria, aerodynamic stability and control parameters, and the performance penalty incurred by aerodynamic drag. The test program will include both force and pressure measurements through the flight Mach number range for which these effects are significant.

Wind tunnel testing of selected shapes at velocities never before studied will be necessary to determine re-entry vehicle characteristics. Particular emphasis will be placed on control surface capability and heating problems. Manoeuvrability limits, g loadings, re-entry corridor characteristics and subsonic landing characteristics must be determined in support of the engineering design program.

Integration and correlation of the ground wind-tunnel testing with the high-speed re-entry flight test program is essential.

4.3.1.1.4 Material Tests

A materials development test program will be undertaken to determine the allowable design strength values and provide design information on the selected structural materials over the appropriate temperature ranges for the base metallic, ablative surfaces, and welded joints. Particular emphasis will be placed on tendency toward brittle fracture under service conditions and in selecting materials for re-entry at 37,000 ft/sec. The testing program will consist of at least the following:

Materials.
a. Smooth and notched static tensile tests of the selected materials.
b. Static tensile tests of welded joints, both fusion- and resistance-welded, for the selected joint configuration for each type of sheet material.
c. Smooth and notched static tensile tests of the selected extrusion and forging materials.
d. Notched impact tests of the extrusion and forging materials.
e. Low-cycle, high stress fatigue tests of welded joints made by the fusion and resistance methods for the selected joint configurations in sheet materials.

This data will be accumulated for the appropriate temperature ranges, i.e., from elevated re-entry temperatures to the cryogenic temperatures in the tanks, as dictated by the projected environmental requirements. In addition, supporting tests such as metallographic examinations and chemical composition determinations will be made as required.

4.3.1.1.5 Static Tests

Static test program will include design load structural substantiation testing to demonstrate structural integrity of the Manned Lunar and Cargo Payloads.

Structural substantiation testing to design loads and temperatures will be accomplished on a full-size stub tank, identical (except for length) to the Lunar Landing Stage tank. This will ensure that load introduction and takeout will be representative of the flight article.

One complete interstage adapter will be tested to ultimate design loads under appropriate environmental conditions. The adapter will be attached to a stub tank identical to the Lunar Launching Stage tank section in every respect except length, to ensure realistic load introduction and takeout.

A stub tank will also be used to demonstrate the integrity of the Lunar Launch Stage tank construction under design loads and environments. Methods of introducing the payload vehicle loads into the adapter section and thus the Lunar Launch Stage tanks will be determined.

Tests will also be run on full-size tank bulkheads. These will be attached to a segment of typical tank structure, adequate to allow the bulkhead behaviour to be representative of that of the flight article under design conditions. These bulkheads will be tested to ultimate design loads to ensure their structural reliability at all points within the flight regime.

Ground handling equipment tests will cover critical fittings and joints for structural substantiation of these items under design conditions.

4.3.2 DESIGN EVALUATION TESTS

Component design evaluation testing is defined here as informal testing conducted by the vehicle contractor, or vendor test labs, for the purpose of basic design evaluation prior to production release, sad to pinpoint critical areas in prototype packages.

Qualification testing is defined as those formal tests performed on flight-type hardware to demonstrate compliance with design specifications. A qualification test plan will be prepared approximately 90 days after engineering go-ahead outlining the qualification test conditions. The qualification tests are to be performed in strict accordance with written and approved detailed test procedures, and witnessed by the Air Force, or an approved representative.

The vehicle contractors' test laboratories will conduct these tests, or subcontract and supervise them at an independent testing agency. Components to be tested will be determined during the engineering design effort.

Controlled environmental conditions will simulate conditions that airborne and ground support equipment are expected to experience during manufacturing, shipping, storage, pre-flight and flight.

Environmental testing conditions will be established based on data already obtained in research and development programs on large rocket-powered vehicles and associated support equipment. Conditions for shipping, storage and handling environmental tests are established in current military and commercial specifications. Subsystem, combined subsystems, and structural evaluation tests will be run in appropriate laboratory faculties to investigate component and subsystem interactions, and to prove out structural designs. Acceptance test procedures will also be developed for use in the factory on deliverable hardware.

A test plan describing the basic conditions and test objectives of each factory systems test, along with the checkout parameters and recorded evaluation data, will be prepared.

A final acceptance test will be required at the time the contractor delivers the vehicle to the Air Force. Test conditions will be as close to the flight conditions as is feasible and safe. All systems will be energised and operated simultaneously.

A final acceptance test evaluation document will be prepared for use by the Air Force and the contractor in determining compliance with test requirements.

Systems acceptance test procedures will be based on all critical parameters required to determine proper functioning of each system in accordance with design specifications and drawings. This will assure a co-ordinated effort of vehicle design, test equipment design, and factory acceptance testing.

4.3.3 FLIGHT TESTING

The Lunex flight test program represents a long and expensive effort leading to the first manned landing on the moon. It requires basic research flights, equipment checkout flights, capability demonstration-flights and finally, the manned and cargo Lunar Expedition flights. This type of effort can only be achieved efficiently and at a minimum cost if the end objective is always clear and the program is designed to meet this objective.

The Lunar Expedition flight test program will provide many side, but important apace capabilities. For example: in April 1965 the first orbital flight capability in a true apace vehicle will be possible; in September 1966, man will make his first flight around the moon in a fully manoeuvrable and recoverable re-entry vehicle; and in August 1967 the first men will land on the moon. Essentially these can all be called test flights, but in each case the system is only at the beginning of its capability instead of being a dead-end item. Each of these capabilities may readily be expanded to provide a military capability if necessary.

The flight test program is summarised on the Lunar Expedition Test Schedule. The following major objectives will be accomplished in the indicated part of the test program.

4.3.3.1 HIGH-SPEED RE-ENTRY FLIGHT TEST
Since present wind-tunnel capabilities are limited to approximately 18,000 ft/sec., it is necessary to perform re-entry flight testing at velocities that range from 25,000 to 45,000 ft/sec. The major objectives of this test program are to:

a. Verify or disprove present theories on basic re-entry techniques as extrapolated to the stated velocity range.
b. Determine problem areas and develop new fundamental theory, numerical procedures and testing techniques where required for this re-entry range.
c. Identify the following:

(1) Items that can be investigated further on a laboratory scale.
(2) Specific laboratory facility requirements.
(3) Additional flight tests that must be performed.

d. Support the engineering design program for the Lunex by providing the above data and special shape testing if required.

4.3.3.2 Lunex RE-ENTRY VEHICLE FLIGHT TEST
The Lunex Re-entry Vehicle will be flight tested by various techniques and in varying environments. Each test will be designed to allow the vehicle to proceed to the next more difficult step. The major testing steps are presented below, with the major test objectives for each step.

a. Prototype Drop Test
Prototype vehicles will be drop tested from a B-52, or equivalent, in both an unmanned and a manned series of tests. Each series will be designed to:

(1) Establish landing characteristics.
(2) Measure inherent subsonic, transonic, and hypersonic stability and control characteristics of the vehicle.
(3) Explore the flight characteristics of the re-entry vehicle in every possible portion of the Mach number spectrum.
(4) Train Lunex crews.

b. Orbital Test
Maximum use will be made of SAINT orbital test information and unmanned and manned flights will be accomplished. These tests will demonstrate:

(1) The capability of the Lunex Re-entry Vehicle to operate in the orbital area.
(2) Re-entry capability at velocities of 25,000 ft/sec.
(3) The manoeuvrability of the re-entry vehicle and its capability to land at a pre-selected earth base.

c. Circumlunar Test
This flight will use the Circumlunar Propulsion Stage and the Lunex Re-entry Vehicle. The major test objectives are:

(1) To send an unmanned and then a manned vehicle around the moon and return to an earth lending at a selected base.
(2) To check out guidance, flight control, guidance, communications and life support sub-systems in a true space environment prior to landing on the lunar surface.
(3) To perform manned reconnaissance of the lunar surface.

d. Lunar Landing and Return
The unmanned vehicle flights will check out the Manned Re-entry Vehicle and related systems to provide a completely automatic system before man first tries the most difficult step in the Lunex program. The major test objectives for these flights will be to:

(1) Check out the Lunar Landing and Lunar Launching Stages.
(2) Check out the Cargo Payload's ability to deliver cargo packages to a preselected site on the lunar surface.
(3) Place three men on the lunar surface so that the initial surface reconnaissance can be accomplished prior to the arrival of the Lunar Expedition.


4.3.3.3 LUNAR LAUNCH STAGE FLIGHT TEST
The Lunar launch Stage will be initially checked under orbital conditions to:

a. Demonstrate space environment operation.
b. Demonstrate engine restart after "soaking" in apace for an extended period.
c. Demonstrate automatic checkout, communications, and remote control capability.

The Lunar Launch Stage will then be flight tested with the complete Manned Lunar Payload for the unmanned and manned Lunar Landing and Return Missions.

4.3.3.4 LUNAR LANDING STAGE FLIGHT TEST

The Lunar Landing Stage will be initially checked out by drop testing. These tests will:

a. Demonstrate landing techniques and the capability of the selected landing system.
b. Evaluate the effects of unexpected terrain variation.
c. Determine the effects of malfunctioning equipment during the landing manoeuvre.
d. Evaluate the effects of engine blast on landing surfaces similar to the predicted lunar surface.

The Lunar Landing Stage will receive its first space evaluation in orbit. The major objectives are:

a. Correlate drop-test data with orbital or space operations.
b. To determine the effects of space environment on the stage.

The first Lunar Landing with the Lunar Landing Stage will be accomplished with a Cargo Package as the payload. When this has been completed a Lunex Re-entry Vehicle will be landed unmanned. The test objectives are to:

a. Demonstrate the feasibility of landing large cargo packages on the lunar surface.
b. Demonstrate the feasibility of automatically landing a "manned vehicle" while unmanned.
c. Provide a man-rated system for the Lunar Expedition.

4.3.3.5 CARGO PACKAGE CONFIGURATION
Various configurations for the Cargo Package of the Lunex Cargo Payload will be tested. The objectives are to:

a. Determine the Cargo Payload aerodynamic characteristics.
b. Demonstrate that the Cargo Packages can be delivered where desired on the lunar surface.

4.3.3.6 ABORT SYSTEM FLIGHT TEST
The fact that a system of this magnitude must possess some measure of "unreliability" is recognised and a "fail safe" abort system is required to insure the survivability of the crew. The test objectives for the Abort System Flight Test Program are to:

a. Demonstrate that crew members in the manned Lunex Re-entry Vehicle can be recovered safely in the event of a malfunction.
b. Demonstrate that the Space launch System is capable of shut-down, or thrust vector change, so that crew abort is possible.

4.3.4 CHECKOUT AND TEST EQUIPMENT

The test equipment will be fully automatic with quantitative readout capability for all critical functions. The Lunex checkout equipment will be the same, or compatible with the Space Launching System checkout equipment. The equipment will be capable of checking out the complete booster and payload system as well as any individual, or isolated component, or subsystem. It shall be fully capable of checkout of any one stage, or the re-entry vehicle, as an isolated unit, and will mate with the stage interface functions and furnish appropriate operational or simulated error inputs to the stage systems.

For the time period of interest, it ii entirely practical to incorporate malfunction prediction capability for preventative maintenance. This will entail a computer function which will accurately control and record the input and output signal values to each system or component. Variations in operation will be recorded and compared to predetermined failure values or characteristics and will forecast the remaining service life of the system under test.

The checkout equipment shall be installed in each blockhouse and it may be used in conjunction with the launch area. This same equipment shall be utilised in the vehicle manufacturing checkout and test functions, as well as in the launch complex, receiving inspection, and maintenance facilities.

The blockhouse equipment will monitor the launch control system commands and inputs as well as those of the payload. Because the launch control equipment will display only go/no-go signals, the checkout and test system will furnish quantitative displays of any function under question for human appraisal and decision.

When the systems are flown unmanned and on the early manned lunar flights it will be necessary to provide automatic checkout where appropriate via a telemetry link. As an example, prior and during lunar launch the checkout procedures will be monitored at the earth control station via the telemetry link.
 

 

4.4 PRODUCTION PLAN


At the present time a detailed Production Plan is not available. However, the present preliminary design study will be completed on 30 June 1961 and the final report to be provided by six independent contractors will include their proposed Production Plan. When the study results have been evaluated a Production Plan for the Lunar Expedition will be prepared .

Several points are apparent at this time and they are presented for completeness in this plan.
 

4.4.1 QUANTITY
Limited quantities of early equipment will be required until the test program improves and increases the capabilities of each item and production quantities became possible. Then, as development and testing proceeds the equipment will become more standardised and production techniques will become more applicable. When the Lunar Landing and Return flights are initiated it will be necessary to launch vehicles at rates that vary from one to two flights per month. When the Lunar Expedition is actually underway the launch rate will remain at a rate of two per month for an extended period. Considering the size, weight, complexity, and importance of these vehicles this represents production rates even when compared to past aircraft or missile production programs.

4.4.2 QUALITY
The inherent reliability of the systems required for the Lunar Expedition program will be maximised by good design practice. Reliability testing represents a major effort of the test program, but the achieved reliability of these systems can only be maintained during production by an excellent quality control program. This means that good organisation, adequate manning and early recognition of the quality control problem is essential. Close co-ordination is required between the quality control personnel and the reliability personnel in the design, development, and test program if the reliability program and the test results are to provide the proper guidance so that quality can be maintained throughout the production effort.

4.4.3 LOCATION
It is anticipated that most of the major systems and sub-systems can be manufactured at facilities, or locations presently in existence and available to the aerospace industry. However, the possibility does exist that certain items, such as the first stage solid propellant stage, may be manufactured at the Lunar Launch Complex due to its size and transportation limitations. These particular items have not been specified at this time, but this will be done as soon as possible.

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