Report by Officer in Scientific Charge - H.G.R. Robinson

Report prepared in conjunction with Mr C.R. Hume, Hawker Siddeley Dynamics - Deputy Officer in Scientific Charge.

F1, a test vehicle for the first stage of the ELDO satellite launcher vehicle Europe I, and the first vehicle to be launched in Phase I of the development program, reached Australia on 18th January, 1964. It was unloaded at Adelaide in the 19th and transported to Woomera at the 24th January.

From the period 26th January to 10th March the vehicle remained in the preparation area, Lake Hart, Woomera. During this period a number of modifications at the vehicle were completed and a preliminary checking of the vehicle undertaken.

Some delay in erecting of the vehicle and the launcher occurred whilst outstanding work at the launch site was completed, end the vehicle was finally erected on Launcher 6A on the 11th March.

After some further delays due to site equipment testing, ground supplies were connected to the vehicle on the 19th March and the final preparation of F1 for launching was commenced on the following day.

The preliminary testing and proving of the vehicle culminated in a successful Triple Transfer test on the 15th and 16th April. The "Triple Transfer" test simulates a firing except that dummy igniters were fitted. Propellents are loaded ant all events, including range participation, are carried out to the preplanned launch time scale, as for an actual firing.

Preparation for the Static Firing commenced en the 20th April and a very satisfactory trial was carried out on the 30th April after delays on the previous day due to a minor plug fault.

A further Triple Transfer teat was conducted on the 12th and 13th May as a joint vehicle/range system training exercise, and preparation for launch on the 25th May commenced.

During the period immediately prior to 25th May outstanding problems concerning range safety and instrumental coverage were resolved with the Range Authorities.

In view of the unsettled weather at Woomera in May - June, it was necessary to take advantage of every chance of fine weather if undue delays in firing were to be avoided. For this reason a decision was taken on the 24th May to continue with preparation for a firing on the 25th, aiming at a firing slot between I and 3 pm, based on a forecast giving a 50% chance of acceptable weather during that period. However, deterioration in the weather during the morning of Monday, 25th May, forced a postponement of the trial after holding the sequence at minus 2 1/2 hours (pre-liquid oxygen filling).

On the basis of further weather forecasts, firing slot allocations for Thursday 28th and Friday 29th May were arranged with the Range. However, during the overnight pre-firing routine checks of 27th - 28th May, a fault was detected in the space gyroscope package. This component was replaced, its alignment checked, and preparation continued for firirng between 1 and 3 pm on Friday, 29th May. The sequence was held at minus 2 1/2 hours (pre-loxfill) until 1.15 p.m. when the decision was made to again postpone the trial due to excessive cloud cover and high winds.

Examining the weather pattern actually experienced during the week 24th - 30th May, it was decided, in conjunction with the Range, to attempt to rescheme the firing slot times to allow some flexibility in choice between morning or afternoon firings, and where possible, to considerably extend the duration of the slot, This was found possible, though it entailed very extensive reorganisation of operations, staff movements, and logistic support; both as regards vehicle and range personnel.

Following battery changes and rechecking of vehicle and range systems, the 3rd firing attempt was made on Tuesday 2nd June, after a weather forecast indicated that suitable weather could be expected during tnt morning. The attempt was aborted at minus 2.6 ascends, after engine light up, at 9.48 a.m. The abort was diagnosed as due to automatic stop action initiated by the Safety Interim Checkout Equipment located in E.C. 6. This equipment automatically detected a momentary failure of one of the four vehicle WREBUS command break up receivers to recive a 'prohibit' signal.

After exhaustive rechecking of the vehicle and ground command break up system, and revaluation of the vehicle following the operation of the engines, the firing day was rescheduled for Friday, June 5th. As as result of weather forecasts, a decision was taken to wait for a long slot, of nominal duration 9 a.m. to 3 p.m. At dawn the weather trends were excellent, and following satisfactory progress of the count down overnight, preparation was continued for a target launch time of 9.15 am.

The vehicle was successfully launched at 9.11am. after an extremely smooth and efficient final count down, both as regards vehicle and range.

Weather conditions were excellent and visibility exceptional.

The vehicle lifted off and programmed downrange according to plan, its flight path and walking impact point following closely to nominal. At about 130 seconds, however, telemetry records indicated the commencement of incipient instability, This became marked at 140 seconds, developing into an uncontrolled corkscrew at 145 seconds. At 147.5 seconds the engine ceased thrusting, some six seconds before the planned time for engine cut. The termination of powered flight has been diagnosed as arising from fuel starvation caused by the maneuvers of the vehicle during its final period of instability. An analogue simulation of the control system related to detailed examination of the telemetered in-flight data has established that the dynamic instability of the vehicle before engine cut arose from a negatively damped fuel slosh mode. Steps are in land to avoid this effect in the next Phase I firing.

Except for somewhat inadequate behaviour camera coverage around the region of engine cut-off, due to the long ranges involved, the records obtained by range instrumentation are excellent, and well up to theoretical expectations.

Telemetry records of good quality were obtained from all three senders throughout powered flight, and high quality data obtained from the majority of optical equipment. Radar tracking was excellent.

The extremely good visibility enabled observers to obtain behaviour data up to apogee, at about 4 minutes, when break-up was observed. This is supported by telemetry and radar data gathered at this time.

This summary was followed by a more detailed report. Of interest is the section about the loss of control of the vehicle near the end of the flight due to fuel sloshing:

A lateral transient was observed at 54 seconds, this has been attributed to wind shear, and was handled satisfactorily by the control system.

A lateral oscillation of the vehicle between 96 and 106 seconds at between 1.9 and 2 c/s growing in amplitude, but suddenly ceasing at the operation of the programmed gain change has also been analysed and found to be due to "rigid body" instability.

The catastrophic instability resulting in premature engine cut commenced at about 130 secs, as an oscillation in both pitch and yaw planes, at about 1.6 c/s. This eventually resulted in saturation of the hydraulic actuator system and loss of roll control. The engines cut some two seconds after the violent uncontrolled motion became readily apparent on behaviour film records. Combined analogue and digital simulation of this instability has verified the time of onset of the phenomena, and shown that the negative damping of the fuel slosh mode becomes increasingly large from 130 seconds to engine cut. As a result of the analyses of the instability H.S.D. and R.A.E. have been able to recommend steps to avoid this effect occurring in the F2 and F3 flights. A reduction of gain, whilst not changing the sign of the damping term in the fuel slosh mode, enables the onset of unacceptable oscillation to be postponed to after the nominal engine cut time of 154 seconds. This analysis is described in more detail in further notes now in preparation.

The lateral motion at engine cut continued throughout free fall up to about 240 seconds from launch, as the vehicle approached apogee. At this time, the re-entry accelerometers, operated by the centripetal acceleration resulting from tumbling were armed according to programme. Evidence from telemetry and visual observers indicates that at this time the break up charges operated, but did not ignite the residual propellent which would be expected to be remote from the central break up charges at this time due to the tumbling motion of the vehicle.

The flight plan may be summarised as follows:

Velocity at engine cut 9625 ft/sec (best information at 20/7/64.)
Height at engine cut38.9 n.m.
Distance downrange 51.8 n.m.
Impact range 548 n.m.
Impact time 850 secs
Apogee height 85 n.m.v
Apogee range 270 n.m.

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