STUDY SUMMARY- "Feasibility Report: Modification of A-12 Vehicle for Air Launched Orbital Reconnaissance System"
Lockheed Aircraft Corporation, 7 September, 1962
RECOMMENDED CONFIGURATION-
1. Launch Vehicle- The basic A-12 proved satisfactory as a launch vehicle configuration. Modifications will
enable the full altitude and speed capability of the A-12 to be used in launching the booster and payload. Using
the maximum A-12 performance at launch minimizes the requirements for booster size. Changes required for
the A-12 are provision for a second crewmember and the structural attachment of rails to the underside of the
A-12 fuselage. This permits structural attachment, checkout, in flight traverse, and separation of the booster
and payload. The existing A-12 or AF-12 subsystems possess the necessary power, cooling, and support
functions required by the booster and payload.
2. Booster system- The booster system selected is based on mods of the Polaris A3 and is described under
Configuration C in LMSC Report SP 2-374. The mods consist of addition of a third booster stage and payload
section, with recovery capsule similar to the current Agena/Discoverer. This system had been designed for
injecting a payload in orbit from a submarine. With careful design, this booster can be attached to the
underside of the A-12 with reasonable mods.
LAUNCH VEHICLE PERFORMANCE- The important performance parameters for the launch system are: radius to
launch area, velocity at booster separation, altitude and attitude at booster separation. The attitude of the
booster is a function of the launch vehicle attitude at separation, the booster system stabilization, and the
time delay between separation and ignition. Normal range of values are:
Cruise out to 3.2M (2 refuels)......................................2,340 n. mi.
Cruise out to 0.8M.......................................................2,340 n. mi.
Attitude at booster separation.....................................20 degrees nose-up
Velocity at booster separation......................................2.5M to 3.0M
Time from Separation to Ignition..................................5 seconds
DESCRIPTION- LAUNCH VEHICLE
The performance factors required for the subject mission required imparting the missile an initial condition
velocity vector, range for the required mission, and navigation and guidance systems which, in conjunction
with the missile systems, provided satisfactory launch parameter accuracy. Studies showed that the cruise
drag of the combined missile/A-12 arrangement allowed adequate range with IFR and provided the required
performance margin for launch. The two configurations studied were: a) missile carried on top of the fuselage
b) missile carried beneath the fuselage.
Considerations affecting the location of the missile stowage (over vs. under) were:
a. Delivery-
1. Performance (velocity vector) at first stage ignition.
2. Safety
b. Take-off Performance-
1. Ground clearance.
2. Nose gear unstick.
c. Cruise Performance-
1. Drag
2. Stability and Control
d. Ground Handling.
OVER-STOWAGE
Stowage on the upper fuselage (fig. 1) is performed by sliding the missile onto a track. Separation
is accomplished by extending a drag chute and pulling the missile aft. Assymetric fuselage loads
require a severe structural beefup to support this concept. Separation safety is considered much
more critical, and additional separation pre-ignition timing degrades performance, compared to
belly launch. The vehicle affects upper surface flow and reduces the predictability of the directional
stability compared to the underside installation. A study was made of the aircraft response during
the launching of a Polaris missile from the track atop the A-12. It was assumed that the missile was
launched by being pulled aft along the rails from atop the A-12. A parachute deployed from the
missile provided the pull. As the missile slides aft, it moves the CG of the A-12 aft until, at
disengagement, the A-12 CG is at 63%. This large instability, about 30% MAC, results in very rapid
pitch motion in response to any pitch disturbance. These motions result in positive or negative load
factors on the A-12 sufficient to break the airplane. Ideally this type of launch could be made at a
zero gee load factor except that this manuever cannot be made precisely enough to eliminate
disturbances that would cause the airplane to diverge. These studies were made using the A-12
simulator and a typical response is attached for a cruise condition launch with the airplane in a
pitch over to 0.25 g. It is apparent this launch procedure is not practical.
PREFERRED ARRANGEMENT- UNDER STOWAGE
The proposed installation is shown in figure 2, General Arrangement- AP-12. The missile is secured
on a track running along the bottom of the fuselage. The 421.6 inch length and 54 inch diameter
permits one foot clearance with the gear and ground plane. Some modifications to the landing gear
will be required to allow full extension length with limited oleo compression during the taxi and
takeoff condition and nose-wheel liftoff rotation. The takeoff CG and gross weight are met by loading
50% fuel, and sequencing this to counterbalance the forward missile position required for takeoff and
gear retraction. The missile tail fairing is aft of the main gear for the takeoff position, with the same
ground clearance angle used as on the A-12. The nose gear retracts forward of the missile, and the
main gear and doors close between the missile and tail fairing, as shown. The tail fairing is moved
aft to cruise position along the fuselage track, and locked. The missile is shifted aft, while the fuel
transfer and stability augment compensate for trim change. The aft position is selected for optimum
cruise longitudinal and directional stability condition. The missile snugs into the fuselage/missile
fairing and the boat tail is positioned into the tail fairing. The above operation is reversible and can
be checked out on the ground by removing the landing gear inner doors. All functions will then be
validated in the cruise position, including seal adjustments, before returning to the takeoff position.
The underneath stowage arrangement allows for ready mating and ground checkout, safe separation
and improved cruise configuration. Performance parameters from the launch to missile ignition are
minimized with the stabilized missile drop procedure.
DESIGN DATA-
The aircraft/missile combination at takeoff weighs 117,000 pounds, or the same as the A-12. 70,000 pounds
of fuel are available after in-flight refueling, at a maximum flight weight of 141,500 pounds. Landing weight
is 52,000 pounds. Weight and balance of the aircraft plus missile, showing the missile shift, indicate an 8%
aft movement with the missile traverse. The airplane is designed for a 2 g normal load factor, which will
account for the mission design conditions.
MISSILE STOWAGE AND LAUNCH PROVISIONS-
a. A group of hooks for missile support (three stations) and integral release to free fall.
b. Fairings for the aircraft/missile fillet and an aerodynamic fairing for the missile boat tail.
c. A track for missile traverse from the ground position to the cruise/launch position. Traverse power is
supplied by a rack and pinion arrangement.
d. A seperable, expendable cooling jacket for missile case cooling under high speed cruise conditions.
e. An extendable leaf tail flare for missile stabilization. This is jettisoned at ignition.
f. Cooling and electrical connections are stowed within the joint fairing lines.
A base plate is attached to the aft end of the basic missile, which engages the tail fairing section mounted aft of the launch vehicle landing gear by means of spring loaded hooks. The flaring mechanism (inverted umbrella) contained in the tail fairing section is spring loaded in the faired or retracted position. A fraction of a second after separation, an electrical signal from the missile is programmed to initiate a pyrotechnic device to deploy the umbrella. Electrical connection is made through the missile base plate. At first motor ignition, the entire tail assembly is jettisoned, including the attached base plate, by means of a pyrotechnic device.
LAUNCH-
1. Launch Parameters- The best target launch condition is a compromise between the carrier aircraft capability in terms of flight path angle, load factor, and Mach number, and missile performance in terms of altitude and payload to orbit. One basic cutoff point is the minimum orbit altitude of 80 nautical miles. Below this altitude, the orbit vehicle performance is degraded by the atmosphere. A second cutoff point is the maximum flight path angle that can be achieved by the carrier at the launch point. This angle decreases with increasing Mach number and is limited by the 2.0 g load factor capability of the aircraft and missile. For example, the maximum flight path angle at Mach 3.2 is about 15 degrees, increasing to more than 20 degrees at Mach 2.5. Conversely, orbit altitude decreases with decreasing flight path angle at launch. A study of conditions based on one post-ignition missile guidance program indicates that, within these limitations, a payload of 900 to 1000 pounds can be placed in an 80 n. mi. orbit for carrier launch speeds between Mach 2.5 and 3.0. These studies are being continued to determine the best post-launch missile guidance program to place the maximum payload into orbit. It is apparent from the analysis which has been done that there is a substantial array of parameters acceptable to both the carrier and the missile which will place payloads of 900 pounds or more into orbits 80 n. mi. or higher.
2. Carrier Manuever and Escape- The launch manuever is primarily a 2 g pullup from a high speed level flight condition. The objective is to obtain the highest flight path angle consistent with normal piloting techniques and flight safety. The escape maneuver after launch is a 2 g rollout from the pullup to increase lateral separation as well as vertical separation between the missile and aircraft. At ignition, the vertical and lateral separation will each be in excess of 500 feet.
3. Separation- Studies of means for stabilizing the aerodynamically unstable Polaris during the launch phase to ignition indicate that a skirt attached to the tail fairing can be expanded upon launch to provide positive static margin of one diameter. This device is seen in figure 3. Alternates considered to provide compensating moment to balance the two caliber unstable vehicle included fins and a drag chute. The necessary drag chute would be approximately 16 feet in diameter and provide 1 g of drag force in order to generate the stabilizing moment. This method degrades launch performance in that 160 feet/second velocity is lost in five seconds of free fall and a three degree flight path angle. This is equivalent to approximately 70 pounds of payload in the design launch regime. Tail fin stabilizing devices, although representing a low drag contribution for balancing moment obtained, introduced stowage problems for the cruise condition and more complex mechanization, assuming fins stowed within the tail fairing. The extendable tail flare results in a decelerative force of 0.3 g during launch and would result in a speed loss during the 5 second missile free fall of 50 feet/second. This is equivalent to 0.05 Mach. The tail fairing and skirt are separated at missile ignition. Vertical separation between the missle and aircraft at ignition will be greater than 500 feet. The flight path angle of the missile at ignition will be 1.5 degree less than the time of launch.
B. Cruise Performance- The AP-12 performance is based directly on an A-12 performance capability corrected for additional weight and drag of the external missile. Missile drag increments are taken from the Polaris wind tunnel data. Mach 3.2 and subsonic mission are shown on figures 4 and 5. Both missions are based on takeoff from a secure inland US base and a missile launch point several hundred miles east of Hawaii. Initial takeoff is with a partial fuel load to keep airplane taxi and takeoff loads at a reasonable level. The airplane climbs to 25,000 feet and is refueled to 141,500 pounds. The second refuel operation occurs over the mid Pacific. Following this operation, the airplane cruises to the launch point and launches the missile at 80,000 feet and Mach 3.1. Fuel allowance is made for a 360 degree turn at the launch altitude and speed. Following launch, the airplane descends toward the alternate field in Hawaii to a rendezvous with a tanker. The return leg to the takeoff point is by supersonic high altitude cruise for both missions. Fuel reserves at mid-ocean refueling point are about 8,000 pounds, and the fuel aboard at the post launch refuel point is sufficient to make an alternate field in Hawaii with a 3,000 pound reserve.
MISSION SUMMARIES
Mach= 3.2 Cruise Out Weight Range Altitude Range Time Increment
Takeoff and climb to start 100,000 4,500- 25,000 ---
refueling over base
Climbout, Cruise, and Descend 141,500- 93,700 25,000-78,000 59 minutes
to Refuel
Climb and Cruise to Launch Point 141,500- 97,500 25,000-80,000 46 minutes
Climb, Cruise, and Descend to 118,000- 74,100 25,000-90,000 95 minutes
to Home Base
Reserve Fuel at home base--- 24,700 pounds
Reserve Fuel at Alternate base-- 3,000 pounds
MACH 0.8 Cruise Out-
Takeoff and Climb to Start 100,000 4,500-25,000 ---
Refueling over Base
Cruise Out to 2nd Refueling 141,500- 94,600 25,000-35,000 158 minutes
Reserve Fuel--- 8,500 lbs.
Cruise to Climb to Launch 141,500- 97,500 25,000- 80,000 107 minutes
Climb, Cruise at M= 3.2 and 118,000- 74,100 25,000- 80,000 95 minutes
Descend to Home Base
Reserve Fuel at Home Base-----24,700 lbs.
Reserve Fuel at Alternate Base---3,000 lbs.
