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:
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:
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.