by Damien Burke
Thrustmeter
Monitoring the performance of the TSR2’s engines would not be simple. The Olympus engines installed in a Vulcan were easy to meter, as power output was basically decided by one simple variable, the amount of fuel flow. On the TSR2 the calculation would be more involved, with the list of variables including engine fuel flow, reheat fuel flow, intake position, reheat nozzle position (and thus area), the amount of flap-blowing bleed and also water flow. To monitor all these for both engines, along with rpm, turbine temperature and so on, would result in thirty or more gauges and warning lights for the crew to deal with. This was not considered practical for a two-man crew with their high workload at low level. Accordingly, this mass of instruments was reduced to a single thrustmeter, along with centralized warning signals for rpm and turbine temperature.
A joint Naval/Air Staff Target (No. 975) had been issued in 1957 for a gross thrust meter (also known as engine pressure ratio (EPR) indicator because it used aircraft static and jetpipe pressures to produce an indication), but by 1960 little to no work had been done to meet this requirement, apart from basic research. This was primarily because the various engine firms believed there was no need for such an instrument when rpm and jetpipe temperature indications were adequate for the task in most circumstances. However, in many cases they were talking of multi-crew aircraft with crew members available to deal with the intricacies of calculating or looking up the correct rpm figures to set up the required flight conditions. (For example, a cruise climb at a specific airspeed given a particular outside air pressure and temperature.) In aircraft such as fighters or the TSR2 itself, reduced crew and cockpit room would make this sort of work a tricky proposition. Nor could the pilot rely on a marked throttle position to give a particular power output, as the relatively small amount of throttle movement available to him would cover a much larger variation in engine power, and so there would inevitably be variations between individual aircraft (or, rather, variations between the throttle lever position and any particular engine’s response).
In contrast, many American aircraft used EPR indicators successfully, particularly to enable checking of adequate thrust before take-off. A ‘gross thrust indicator’ was proposed by BSEL for the Olympus 22R/ TSR2 project, and as such an instrument was also required for the NA.39 (ascertaining take-off thrust before a catapult launch was a critical requirement), development of a thrustmeter became a Category 1 item. The difficulty of this apparently straightforward task was a good indicator of the unexpected complexity lying in wait for the TSR2’s creators around almost every corner.
Elliot Brothers was contracted to develop the thrustmeter, but it was not until August 1961 that a firm development plan was in place. Both BSEL and BAC had put together proposals for how such a meter would measure, calculate and display its readings. Two stages of thrustmeter would be produced. The Stage 1 version, based on BSEL’s specifications, was to be fitted to the first five development-batch aircraft, and the slightly less accurate but more logically presented Stage 2 version as envisaged by BAC was to be fitted from aircraft No. 6 onwards. The Stage 1 meter suffered from a display that did not take into account ambient conditions, so on a cold day in full reheat it would indicate greater than 100 per cent. Similarly, on a hot day, it would indicate less than 100 per cent. In a way this was indeed accurate, as temperature did affect performance, but BAC’s Stage 2 proposal would present 100 per cent in full reheat regardless of outside air temperature, and this was a more logical presentation of the measure of available power versus power output.
Fire protection
The TSR2 as built was effectively a flying bomb, with fuel tanks surrounding the hot engines and weapons bay, and no blade containment built into the engines. Normal RAF practice was to compartmentalize engine bays into numerous fire zones, isolating particularly risky areas and adding fire detection and suppression systems where possible. On the TSR2, BAC and BSEL believed that the high temperatures reached by the entire engine and its casing made compartmentalization effectively pointless, and while the engine installation was divided into zones 1, 2 and 3, conventional fire zoning was not practical. The result was that zone 1 enclosed the entire engine and jet-pipe installation, zone 2 was the annular space between the titanium engine heat shields and the engine tunnel walls (the fuel tank walls), and zone 3 was a small area enclosing the bay under the reheat pipe that housed the fuel pipes to the reheat pipe manifold and the nozzle control trim unit. Access to zones 1 and 3 for firefighters’ hoses was provided by fire panels with break-in circular panels within them.
Engine installation fire zones. BAE Systems via Brooklands Museum
The RAF research department dealing with fire disagreed vehemently with the lack of conventional fire zoning, and the result was a number of requirements to reduce the likelihood of an in-flight fire. Firstly, the hydraulic fluid to be used was DP.47, also known as Silcodyne H, a supposedly low-flammability fluid designed for high-temperature areas. Secondly, fuel pipes would be double-walled (later this was extended to oil and hydraulic pipes too), despite strenuous objections from BAC. Double-walled pipes were heavier, took up more room and could actually increase the chances of leakage. The result was just as BAC feared; the double-walled pipes were a total nightmare, and the company ended up getting a concession to revert to normal pipes later in the programme.
Both zones 1 and 3 would be cooled and ventilated by ram-air bled from the air intakes via the tunnel cooling valve between the intake and engine lip. This airflow would serve several purposes: cooling, removing possible explosive gasses (such as a fine misting fuel leak), and also distributing methyl bromide extinguisher spray throughout the engine tunnels if the fire extinguishers were activated. Methyl bromide, a toxic chemical widely used as a pesticide, fell out of favour for fire extinguishing purposes in the late 1960s and was finally banned over thirty years later. This airflow would exhaust around the nozzle shroud (relieving some of the afterbody suction in the process and contributing to drag reduction). During ground running or at low speeds there was a danger this flow could reverse, so in these cases the tunnel cooling valve would be closed and ram air directed into zone 1 by ground-running doors on the fuselage sides below the wing root. As with so many items on the aircraft, these doors had a backup. If the electrical motor that opened them failed, they had inset spring-loaded doors which would be sucked open as the pressure differential rose.
Engine tunnel cooling arrangement. BAE Systems via Warton Heritage Group
Engine tunnel cooling details. BAE Systems Heritage via RAF Museum
The layout of the fire extinguishing system. BAE Systems via Warton Heritage Group
Fire detection was by means of Graviner firewires running throughout the accessories and reheat control bays, and also in rings around the engine. High temperatures would alter the resistance and capacitance of the wires and trigger warnings in the crew’s cockpits on the central warning panel (CWP): ENGINE 1 and 2 lights for zone 1, and REHEAT 1 and 2 lights for zone 3. Pressing the appropriate button would fire the relevant extinguishers; these would also be fired automatically by inertia switches in the event of a crash landing.
Engine test failures and the Vulcan flying test bed
Testing at the NGTE showed up a variety of minor faults with the test engines, and resulted in various design changes (such as the switch to titanium blades mentioned previously). However, of more concern were the catastrophic failures, of which there had been several by December 1962. Two were of particular concern. On 3 November 1961 engine No. 2203 was undergoing an endurance cell test when a flame tube wall broke up owing to fatigue. Consequently the LP driveshaft overheated and fractured, resulting in the overspeed and ejection of an LP compressor disc. Improvements were made, resulting in a single-piece flame wall being introduced. Then, on 16 May 1962, engine No. 2206 suffered a similar failure during another endurance test. An LP compressor disc overheated owing to inadequate cooling and burst, scattering shrapnel through the eng
ine, though on this occasion the disc was contained.
Bristol Siddeley Engines had requested a Vulcan B.2 airframe for use as an FTB, which could then be used for both the 22R and 21 engines. The protracted contractual negotiations on the 22R project, however, soon meant that combining the two projects in a single aircraft was not going to be possible. Making further Vulcan B.2s available would have delayed that much-needed upgrade to the V-force, so the Vulcan provided by the MoS was B.1 XA894. Work began in June 1960 to modify it appropriately. This was complicated by the relatively poor state of the airframe, which required a substantial number of modifications to bring it up to a basic standard even before work began on modifying it for use as a test bed.
The 22R’s reheat jetpipe made it considerably larger than the earlier marks, so it could not replace one of the Vulcan’s existing engines. Furthermore, the value of the FTB would have been seriously reduced by having to use a specially modified 22R mounted in a completely different manner to the engines destined for the TSR2. Accordingly a ventral installation was chosen, with the engine mounted in a fairing that, internally at least, closely resembled the area of the TSR2 airframe into which the engine would be installed. As the Vulcan airframe was decidedly subsonic there was little value to be had from including a representative TSR2 intake, and there was also the problem of the Vulcan’s nose undercarriage leg. This could possibly throw debris into a centrally mounted intake during take-off and landing, so the intake needed to be ahead of the nose gear leg, and to avoid interfering with the leg’s operation, a bifurcated intake was necessary.
The LP drive shaft from Olympus 320 No. 2203, showing the ruptured area; compare with the undamaged shaft in the background. Rolls-Royce Heritage Trust
The complete unit included not just the basic engine but also the TSR2 gearbox, complete with Plessey CSD unit, starter, alternator and two TSR2 fueldraulic pumps. Because of the fuel-flow requirements of the engine when in reheat, two additional 520gal (2,364L) fuel tanks and pumps were installed in the Vulcan’s bomb bay. These were, however, still fed by the Vulcan’s existing fuel system, and could not be replenished at the same rate as fuel was used by the Olympus 320. This introduced a limit of 10min in reheat within a 30min flight time. A section of lagging protected the underside of the Vulcan in the vicinity of the 320’s reheat pipe, and a periscope afforded the crew some view of the entire installation (and was intended to be used for cine camera coverage, though no useful results were ever gained from it). The aircraft was expected to make its first flight with the Olympus 22R fitted at the beginning of August 1961, but delays with both the aircraft fit and the engine test runs put paid to this plan.
During December 1961 a successful 24hr endurance test run of an Olympus 320 was made in one of the Pyestock test cells, resulting in flight clearance for the FTB installation. After the overheats that had resulted in catastrophic engine failure in the test cell, a temperature gauge for the flight test observer’s position on the Vulcan FTB had been introduced so that an eye could be kept on the LP driveshaft temperature and the engine shut down if it began to overheat in this area.
The first ground run of the installed Olympus 320 was made on 31 January 1962, and the first flight followed on 23 February. After 46hr 1min of ground running and in-flight running, the engine and reheat unit were removed and replaced by improved examples, and reheat flights began in June. A further 41hr 45min of running was carried out (including a noisy appearance at that year’s Society of British Aircraft Constructors’ air show at Farnborough) until 3 December 1962, when events took a dramatic turn for the worse. The vulnerable LP driveshaft was about to demonstrate its destructive power once more.
XA894 had been positioned on the engine detuner stand at BSEL’s Filton plant for a full-power ground run with reheat. Observers outside the aircraft saw an orange flash extending from the aircraft’s belly to above the fuselage, and the entire airframe visibly lurched forward. In the cockpit the test crew felt a massive shock and saw the flash from the explosion underneath the aircraft lighting up the gloom outside. With all the fire warning lights glowing, the crew immediately abandoned the aircraft while the attending fire crew, in their particularly shiny and new fire tender, responded with commendable speed and began spraying foam over the burning aircraft. However, fuel was by now pouring out of XA894’s ruptured fuel tanks, and was running down the gentle incline towards the fire tender. Within minutes the fire crew’s position was untenable, and they had to abandon their efforts to fight the aircraft fire and concentrate on saving their own tender. This was to no avail, as the supply of burning fuel was seemingly inexhaustible. Soon they were forced to retire to a safe distance and leave both their tender and XA894 to burn through the night.
Avro Vulcan B.1 XA894, showing the ventral Olympus 320 installation with Y-shaped intake to clear the nose gear leg. Brooklands Museum
XA894 made a spectacularly noisy appearance at the 1962 Farnborough air show. Nowadays, alas, the gentlemen of the press are not permitted to stand quite so close to the runway. Brooklands Museum
A BSEL advertisement of October 1962. Rolls-Royce Heritage Trust via Brooklands Museum
With its Olympus 320 in reheat, XA894 was quite capable of flying along with its own engines at idle. Brooklands Museum
The cause of the failure was once again a rupture of the LP compressor driveshaft. The second-stage LP turbine disc was ejected from the engine and exited upwards, carving a path through the bomb-bay fuel tanks that fed the Olympus 22R unit before being directed back at the ground, where the individual blades were torn from the disc and sprayed through the port wing. Here they ruptured the Vulcan’s own wing fuel tanks and tore away a section of the wing leading edge. What remained of the disc skipped across the airfield, finally coming to rest just metres away from one of the (also shiny and new) Bristol Type 188 high-speed research aircraft.
Subsequent inspection of what remained of the engine was complicated by the intensity of the fire that had taken hold, much of the evidence of the failure having been destroyed. Of the Vulcan itself, nothing was left but the outermost section of the port wing and a large collection of melted alloy and ash, with a few more solid parts scattered here and there. Several months of investigation followed, and there was some evidence of fatigue cracking on parts of the LP shaft around the bearing oil drain holes. Vibration was suspected as the root cause, so BSEL’s initial response was to strengthen the shaft by making it thicker, and also to try to detune the shaft by adding a pair of damping rings that were shrunk on to it. Engine No. 2203 (repaired after its earlier failure) incorporated the improved LP shaft, and bench tests began once again. Thousands of hours of running would be necessary to get reliable results, so a final verdict on the improvement was not expected until the end of March 1965, well over a year after the TSR2 was expected to have flown! Replacement of the Vulcan FTB also became a sticking point, with the MoA paralyzed by indecision. If it was replaced and another engine blew up, perhaps in the air, it would be a disaster from which the programme would be unlikely to recover. However, if it was not replaced, the overall engine programme could be delayed even further. Finally the engine test facilities at Pyestock came to the rescue, and a replacement Vulcan FTB was never authorized.
Disaster strikes; XA894’s Olympus 320 tears itself apart, starting the fierce fire that destroyed the aircraft. Discarded fire hoses litter the foreground of this shot. Rolls-Royce Heritage Trust
Another explosion rocks XA894, its rear fuselage and fin having collapsed by this point. An uncontained engine failure could have had similar results in a TSR2 airframe, in which fuel tanks surrounded the engine installation. Rolls-Royce Heritage Trust
The Olympus 320’s troubled development continued throughout 1963/64, development costs to CA release now having risen to £46 million, nearly six times the original estimates. Despite numerous improvements, fuel consumption was still 10 per cent above specification, which would have a direct impact on the aircraft’s combat radius
. The engine was also proving to be overly sensitive to intake airflow distribution conditions. While BSEL was confident it could improve fuel consumption, it was becoming clear that this was going to take a great deal of extra development, and would still be unlikely to meet the requirement, it being suggested that it would end up falling short by 5 per cent.
Performance at altitude was also going to be a problem. It would be fair to say that nobody, not BAC, BSEL or the MoA, had taken the high-altitude requirements very seriously. It had therefore been a shock when the RAF pressed the case in 1961 to require ‘continuous operation’ at 56,000ft (17,000m). This was a problem, as, based on the engine’s current performance figures, the subsonic ceiling would be 36,000ft (11,000m) (42,000ft (12,800m) with reheat), and above this the aircraft would have to be supersonic, requiring constant use of reheat. But BSEL had been working to a specification that meant the engine need only operate in reheat for 15min periods, and thus the aircraft would only be able to exceed 36,000ft for brief periods. To make matters worse, reheat could only be engaged when the engine was at maximum dry power, and it was limited to 12hr of such use before overhaul was required. At least this limitation could be dealt with fairly easily. The use of reheat was governed by a temperature datum, and an additional datum value could be introduced to permit reheat at less than fully dry power, choosing an appropriate temperature datum to match whatever continuous power setting would be most suitable. The use of reheat in maximum dry power would still be needed for short take-offs and emergencies, however, so a ‘double datum’ configuration would be needed, with the pilot able to select between them.
