The HAWK Story - Royal Aeronautical Society · PDF fileThe HAWK Story Harry Fraser ... apparent that the aircraft then used for the Royal Air Force’s fast jet pilot training were

Journal of Aeronautical History Paper No. 2013/01

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The HAWK Story

Harry Fraser-Mitchell Formerly British Aerospace Ltd

FOREWORD On 10th October 2011, a joint presentation on the Design, Development and Future Prospects for the HSA / BAE “Hawk” aircraft was made to an audience in the Lecture Theatre of the Royal Aeronautical Society, under the auspices of the Historical Group. The speakers were the author, dealing with Design and Development, Mr C. Roberts, Project Pilot, on the T-45 for the US Navy, and Dr A. Bradley, the current Chief Engineer, Hawk, on the Present and Future Prospects. With only a total of just over an hour for the whole presentation, it was impossible for any of the speakers to go into any detail, and the Author felt that it was desirable for the whole story to be written up as a paper, in three parts as presented by the speakers above. He has attempted to do this himself, but relying heavily on data provided by the other speakers and other sources for Parts 2 and 3. It is to be hoped that a future issue might encompass further information from the other contributors to “The Hawk Story”. This paper consists of three parts. PART 1 Starting from the initial investigations by HSA in 1968, the evolution of the HS 1182 project is shown, eventually becoming the Hawk T.Mk.1. The development of the export and strike versions, series 50, 60, 100 and 200 are covered. The U.S. Navy T-45A is only briefly mentioned here – it is covered more fully in Part 2. The evolution of the Rolls-Royce Turbomeca RT 172 Adour is outlined with the help of RR-TM documentation. The Author’s opinions as to why the Hawk has been so successful are given towards the end of this part. PART 2 This covers the adaptation and development of the basic Hawk airframe for the use of the U.S. Navy for training and carrier qualifications, in collaboration with the McDonnell Douglas Corporation, St Louis (originally with Douglas Aircraft Co, Long Beach), both now incorporated into The Boeing Aircraft Corporation.


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The competition, initial developments and the critical design drivers of the modifications eventually agreed are all covered in some detail. Brief mention is made of the lawsuit brought by the US Navy to determine who should be responsible for the extra costs incurred by the need for extra modifications, deemed necessary to meet the requirements of the Specification. PART 3 Since 1995, the approximate time covered by Part 1, the Hawk design has been greatly advanced in many respects, resulting in a large and healthy sales ledger, with new, updated systems and powerplant improvements. Some details of these developments are given in this Part, relying on data provided by the original Speaker, and other published sources. This is not the end of the Hawk story, as further avenues are being actively explored.


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PART 1 DESIGN AND DEVELOPMENT SUMMARY Starting in 1968 with early feasibility studies into what the RAF might need to replace the aircraft in their fast jet pilot training programme, the HS 1182 aircraft was defined. When the Air Staff Target became known, further refinement took place, and the resulting design was selected. Some details of the development of the aircraft’s aerodynamics, structure and systems are described and a few of the problems that arose in the flight testing and their subsequent solutions are briefly dealt with. It was always intended that the Hawk should have an appeal as a light strike aircraft for the export market and the development of the Mk.50, 60, and 100 series is covered as well as the single seat attack version, the Mk.200 series. One section deals with the parallel development of the Rolls-Royce Adour, the engine chosen for the aircraft. 1. INITIAL STUDIES 1.1 Establishment of the requirements In the mid-Sixties, the Air Staff was already thinking of updating the RAF pilot training programme, and in 1964 issued Air Staff Target (AST) 362 for a Gnat replacement. It called for a twin-engine, two seat advanced trainer capable of dash speeds of up to 1.5 Mach number, something like the USAF T-38 aircraft. International collaboration was the ‘flavour of the month’ and the Breguet 121 airframe seemed to be a suitable basis for collaborative study. However, as it evolved, it became clear that it was going to be an expensive trainer, with twin reheated engines, and the drag was such that it even required partial reheat in the approach. But it did look like a candidate for an attack aircraft – as it eventually became, as the SEPECAT Jaguar, the majority of which were single-seaters. This left the RAF trainer programme unfulfilled and in the late sixties it was becoming apparent that the aircraft then used for the Royal Air Force’s fast jet pilot training were increasingly expensive to fly and maintain, and would need to be replaced in the fairly near future. In particular, the Folland Gnat Trainer and the Hawker Hunter two-seater were well into the second half of their service lives. Thus in 1968, on the basis of discussions by Gordon Hodson with Gnat operators, the Future Projects Office of the Kingston-on-Thames works of Hawker Siddeley Aviation started to investigate the requirements for a suitable replacement aircraft, preferably to combine the duties of advanced flying and weapon training.


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K. Gordon Hodson, then in the Kingston office, had seen RAF service and had also been closely associated with the Gnat at Follands. Backed by the Assistant Chief Engineer, R. B. Marsh, he embarked on a series of liaison visits to RAF Training Establishments, with a view to finding out what they would ideally like to see in a new trainer aircraft, and the duties it should perform.

He found that features that were highly desired were (not in any order of priority) as follows:

Low acquisition and life cycle costs

Low fuel consumption and a wide speed range up to high subsonic.

High reliability and hence high utilisation and low maintenance cost.

High structural integrity and from the point of view of the Company:

Low risk.

Ease of manufacture

Simple design.

Development potential.

Export considerations were crucial to cover costs. As a result of the investigations, an internal brief specification was drawn up (see Section 1.7), and the Project Office went to work producing a series of feasibility studies, covering a wide spectrum of types using single and twin engines, tandem and side-by-side seating, straight and swept wings, low, mid or shoulder mounted. A few of these are illustrated in Figures 1 to 8. These are only a selection of perhaps 20 layouts that were studied and assessed. The table below summarises the features of the illustrated configurations.

Figure Type Wing position Cockpit Intake U/C

Mounting Powerplant

1 1182-1 Unswept, high Tandem Wing root Fuselage 1 x Adour

2 1182-2 Unswept, low Tandem High Wing 1 x Adour

3 1182-4 Unswept, mid Tandem Wing root Wing 1 x Adour

4 1182-7 Swept, high Tandem Low Fuselage 1 x Adour

5 1182-8 Swept, low Side / Side High Wing 1 x Adour

6 1182 - Swept, low Tandem Pods Wing 2 x BS 153/ Larzac

7 1182 - Swept, low Side / Side Pods Wing 2 x BS 153/ Larzac

8 1182 - Swept, high Side / Side Wing root Fuselage 2 x BS 153/ Larzac

Notes: All had fixed tailplane with elevators and balanced manual controls.

Unswept wings were rejected as being unlikely to exceed M = 0.8


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Notes: The tandem cockpit was judged to have better vision and lower frontal area, and hence cont. less drag, than the side-by-side cockpit

The high wing position with a low tail was thought to be more favourable at high speed

The low wing position was potentially easier to service and equip with stores

A fuselage-mounted undercarriage (necessary with a high wing) gave an undesirably narrow track

The low intake had a potential ingestion problem, whereas the high one should be clear Twin engines were awkward to install for low drag. There was potential for high speed problems due to interference between podded engines, and a high tail would be necessary. The conclusion reached was to have a swept wing, either high or low (to be investigated), a tandem cockpit, high intake and a single engine. At the same time, studies were made of methods of cost estimation based on past experience – these were fed back to the design people to guide them in offering schemes having significant savings in cost. Over the next two years, all these studies were refined and assessed. One can recall one famous and lengthy meeting in the Project Office, chaired by the Chief Future Projects Engineer, J. E. Allen, when some 17 competing designs, each with its own advocates, were whittled down to one, plus a few variants. This was dubbed the HS 1182 (with later variants 1182 A, 1182 V and 1182 AJ).

Figure 1 HS 1182 – 1 Unswept high wing, wing root intake, fuselage mounted u/c, tandem cockpit, 1 x Adour engine


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Figure 2 HS 1182-2 Unswept low wing, high intakes, wing mounted u/c, tandem cockpit, 1 x Adour engine

Figure 3 HS 1182 – 4 Unswept mid wing, wing root intakes, wing mounted u/c, tandem cockpit, 1 x Adour engine.


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Figure 4 HS 1182 -7 Swept high wing, low intake, fuselage mounted u/c, tandem cockpit, 1 x Adour engine

Figure 5 HS 1182 – 8 Swept low wing, high intakes, inboard mounted u/c, side-by-side cockpit, 1 x Adour engine


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Figure 6 HS 1182 Swept low wing, podded engines, inboard mounted u/c, tandem cockpit, 2 x BS 153 or Larzac

Figure 7 HS 1182 Swept low wing, podded engines, inboard u/c, side-by-side cockpit, 2 x BS 153 engines.


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Figure 8 HS 1182 – 7 Swept high wing, underwing engines, inboard u/c, side-by-side cockpit, 2 x BS 153 engines

1.2 Choice of engines. Different types of engine were also investigated, as shown in Tables 1 and 2. These were appropriate to both single and twin installations. Considering the twins first, it was observed that the J-85 had relatively high fuel consumption. The JT-15 was much better in this respect, but because it had a higher by-pass ratio, its performance suffered at height and speed (Table 2). The Larzac was good all round but was committed to the Alpha Jet and there was some doubt of its availability and of its support from a relatively small company. The argument for the choice of a single engine over a twin was well exercised. For twins, it was said that there must be a better degree of safety in that an engine failure was unlikely to be critical. Against this it was argued that the reliability of jet engines, particularly for those well established, was very good anyway. Furthermore, more training would be required to introduce a student to coping with the deliberate shutting down of one engine to simulate failure. Wryly, some multi-engine training pilots remarked that it was not unknown for students to shut down the good engine when faced with the situation, often with disastrous results.


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Table 1 Comparison of engines at Maximum Take-off Rating, Sea Level, ISA, Static Engine Type Thrust SFC Fuel Flow Air Mass Net Dry (lb) (lb/hr/lb) (lb/hr) Flow Weight (lb/sec) (lb) Twin Engines

GE 85 – 13 5440 1.03 5603 88 1140

JT 15 D – 3 5760 0.525 3025 187 1258

Larzac – 04 5940 0.705 4190 122 1170

Single Engines

Avon Mk.121/122 7575 0.98 7420 123 2502

Orpheus 101 4400 1.061 4670 83 920

RB 199 8394 0.603 5062 156 1160

M 45 H 7760 0.451 3500 233 1500

Viper 600 4000 0.97 3880 58 825

Viper 21 F 4800 0.758 3640 108 925

Adour RT 172-06 5000 0.691 3455 93 1162

Walter Titan/A1 – 25 4000 0.62 2480 Table 2 Comparison of engines at Maximum Continuous Rating, 30000ft, ISA, 0.8 Mach

Net Thrust SFC Fuel Flow lb lb/hr/lb lb/hr

Twin Engines

GE 85 – 13 2160 1.285 2775

JT 15 D – 3 1741 0.846 1472

Larzac – 04 2000 0.973 1946

Single Engines

Avon Mk.121/122 3200 1.19 3810

Orpheus 101 1550 1.315 2040

RB 199 2954 0.835 2473

M 45 H 2240 0.77 1722

Viper 600 1425 1.155 1642

Viper 21F n/a n/a n/a

Adour RT 172 – 06 1770 0.942 1667

Walter Titan/A1 – 25 n/a n/a n/a


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On the design side, the doubling up of engine instruments and engine-related systems meant more space was required and hence more cost and more maintenance hours per flight hour. It was also apparent that two engines of the same total thrust as a single were more expensive. The ‘single engine’ lobby maintained that statistics showed that the number of casualties due to engine problems in single engine jet aircraft was no higher than those for twins – although the ‘twin’ proponents did not accept this, and produced their own statistics to prove it! But we all know about selective statistics! In the event it was decided to go for a single engine. Probably very pertinent to this decision was the fact that Hawkers had never produced a twin-engine aircraft. Looking then at the singles, bearing in mind that something like 5,000 lb SLST (sea level static thrust) was needed and about 1,700 lb in the cruise, the larger engines ruled themselves out almost at once. Considering the engines in turn:

Avon – too heavy, older design, too much thrust, thirsty.

Orpheus – not enough thrust, old design, high SFC.

RB 199 – much too much thrust, and large. Good SFC and modern design; costly?

M45H – bulky but with a good SFC, though the high airflow might be a problem.

Viper 600 – not enough thrust, and high SFC. Aircraft would have to be reduced in size and capability.

Adour 172-06 – Fairly high first cost (but probably Government furnished), a bit heavy, but the reheated version was well established. 50% French origin.

It was decided to go for the Adour, on the promise that there was considerable stretch likely to become available. It was a “modular” engine, and of modern design. It had a reasonably good SFC and was well backed up by Rolls-Royce. Nevertheless, the Viper 600 was kept in mind as a fall-back, and a good deal of work was done on a reduced size of aircraft to assess its capabilities, as undoubtedly it would be cheaper to buy, though limited in its weapon-carrying ability. 1.3 Wing position. Again, there were advocates who favoured a shoulder wing position over a mid or low wing position. The mid wing was a non-starter as the wing structure would pass through the proposed engine position. The shoulder wing gave promise of a better aerodynamic junction, but drawbacks were the resulting narrow track undercarriage, housed in the fuselage (disliked on the Gnat), and the need for a crane if the (one-piece) wing had to be removed for engine removal. However it would be easy to have a low tail, which was probably good from the high speed point of view.


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Wing mounted stores would need a fair bit of lifting for loading, though the wing would probably have to have anhedral. This might result in low wing tip clearance when rolling on the ground. The high intake should ensure that spray, stones etc. thrown up by the nosewheel should not be ingested by the intake. With flaps alongside the fuselage when deflected, there should be good low speed lift. The low wing position was likely to be more difficult from the aerodynamic point of view, in particular the region between the bottom of the intake and the top of the wing, at high speed. There would be a space between the inner edges of the flap which would probably cause some lift loss. It might be difficult to get the tailplane as low as desired. On the other hand, a wide track undercarriage (praised on the Hunter) could be accommodated in the wing, and loading stores should be easier. The wing itself could more easily be lowered for removal, if required. The latter arguments won the case for the low wing, though the aerodynamicists were much concerned about possible problems. They were comforted to some extent by the promise of running early wind tunnel tests at both low speed and high speed, to throw up any possible problems and, hopefully, show the solutions. Thus, work was put in hand to manufacture a half scale half model to be tested with alternative wing positions at low speed in the V/STOL wind tunnel at Hatfield (Figure 9) and a 1/30th scale model to be tested at high speeds at Brough.

Figure 9 Half scale half model for the V/STOL wind tunnel, Hatfield Models of the wing aerofoil section (K14/1 and K14/2) were tested at high speed in the ARA Two-Dimensional wind tunnel at Bedford.


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A 1/10th scale complete model (Figure 10) was manufactured to be tested at low speed at Woodford to supply aerodynamic characteristics.

Figure 10 1/10 th scale first complete model for the Low Speed wind tunnel at Woodford 1.4 The HS 1182 (Document HSK 27, October 1970) This design, shown in Figures 11 and 12, was studied in great detail and was the basic design for the project. Notice the ‘straight’ fuselage in side view, the ‘Teddy Bear’s Ears’ high intakes and the high, flat, all-moving tail. Both tailplane and ailerons were now fitted with duplicated, irreversible hydraulically powered control units on account of the higher speeds.

Figure 11 HS 1182 Trainer (HSK 27)


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Figure 12 Artist’s Impression HS 1182 (HSK 27) Leading characteristics of the HS 1182 (HSK 27) were as follows: Wing span 9,894 mm (32 ft 5.5 in) Wing gross area 18.51 m 2 (199.24 ft 2 )

Aspect ratio 5.287 Taper ratio 0.34

Leading edge sweep 23.8 deg

t/c ratio centreline 11.04 % t/c ratio tip 9.07 %

Fuselage length 12,099 mm (39 ft 8 in)

Tailplane gross area 3.72 m 2 (40 ft 2 ) Tail arm 4,532 mm (14 ft 10 in)

Tailplane aspect ratio 3.55 Taper ratio 0.413

Leading edge sweep 34 deg t/c ratio 8.5 %

Fin net area 3.135 m 2 (33.75 ft 2 ) Fin arm 3,688 mm (12 ft 1 in)

Basic mass (Weapon Trainer, 2 wing pylons)

3,477 kg (7,665 lb) Take-off mass 5,613 kg (12,375 lb)

The structure was stressed to an ultimate load factor of 12 g using a maximum tensile stress not exceeding 45,000 lb/sq.in. The flight envelope showed 8 g maximum normal acceleration with a design dive CAS (calibrated air speed) of 550 knots, or Mach 0.9. The safe fatigue life was estimated as 6000 hours, using a fairly stringent fatigue spectrum. A mock-up of the cockpit had been designed and built, and modified as required to give the best possible layout, investigating internal reflections and lighting. The windscreen and canopy were designed to remain safe following a 2lb bird strike at 450 knots, and a good view was obtained:

Front seat Centreline, 15 o

downward; 40 o

port or starboard, 35 o

downward

Rear seat Centreline, 7 o downward; 40

o port or starboard, 30

o downward


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Folland F4 GT2 Mk.2 seats were specified; these gave safe ejection down to 90 knots, ground level. A very comprehensive equipment fit was offered including a 3,000 psi hydraulic system serving duplicated irreversible powered flying controls for the slab tailplane and for ailerons, (but with a manual rudder), all with push-pull control rods, large double-slotted flaps, air brake and anti-skid wheel brakes. Undercarriage lowering and retraction was also hydraulic and nosewheel steering was offered, but was noted as being potentially costly. A ram-air turbine was proposed to drive the flying controls in the event of major failure. Avionic equipment included VHF, UHF and standby radios, IFF/SSR, TACAN, VOR and ILS. A 5 litre LOX system was proposed and electrical power was provided via a 24v, 6kw engine mounted DC generator. For training purposes, a 7.62mm machine gun was installed in the wing root. As shown above and in the figures, the wing was of a moderate sweep, the Mach number normal to the leading edge being 0.78 at a flight Mach number of 0.85, the expected drag rise point. The aerofoil design was based on experience with the Harrier amongst other research sections, and could be described as having a moderate “peaky” pressure distribution, with a short ‘rooftop’. The aerofoil sections were tested at high speed in the ARA Two-Dimensional wind tunnel, as related earlier. The half model, tested at Hatfield, showed a promising high lift performance with C L max in the region of 1.64 with full flap and undercarriage down. There was little to choose between the high and low wing positions at low speed. The all-moving tail was placed well aft of the fin, so as not to blanket the rudder during an erect spin. The proposed powerplant was the Rolls-Royce Turbomeca RT172-06 having a static sea level ISA thrust of 5,000 lb. It had a good SFC due to the bypass ratio of approximately 0.8, but it was noted that it was expensive and heavy. The high intake incorporated a fuselage boundary layer bleed, similar to that of the Hunter. Full performance data were given, including field performance, cruise and sustained turn performance. Also included in the document was a preliminary work programme, which, assuming a go-ahead in January 1971 and specification agreed in March, showed the first aircraft in June 1973 (18 months) with flight testing, estimated at 600 hours using 4 aircraft, complete sometime in 1975. With full production ordered in August 1973, C. A. Release was expected in March 1976, and a production rate building up to 4 per month for 175 aircraft in total. The programme included static strength tests on the airframe and components, resonance tests, fatigue tests and tests on the pressure cabin. It was proposed to carry out load measurements in flight, for the fin, the tailplane, and various hinge moments.


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The following wind tunnel models were proposed:

Low speed - half scale half model

1/10th scale complete model

1/3 scale intake model

High speed – 1/30th scale and

1/6th scale models

Wing aerofoil section models

Spin – Models for the spin tunnel at Lille University

Free fall spin models. The Air Staff issued a new ASR (No.397) for a trainer aircraft in 1970, with no requirement for sustained supersonic flight, but with a fairly stringent list of other requirements. The requirements of ASR 397 (2nd draft) were compared with the estimated ability of the HS 1182, and unfortunately showed up some discrepancies*, though not severe ones, most being met. For example:-

Requirement ASR 397 HS 1182

Operating speed M = 0.85 M = 0.83 *

Max. level speed 420 kn 500 kn

Optimum Cruise M = 0.6 M = 0.7

Approach speed 60% fuel 120-140 kn 135 kn

Threshold speed 60% fuel 100-110 kn 112 kn *

Time to 30,000ft from brakes off 7 min. 8 min *

Ceiling 40,000 ft 45,000 ft

Field length 4,000 ft met

(Critical) Sortie “E” 1 hour 50min. *

Cross wind 25kn Attainable

Turning performance S.L. 4½ g at 350 kn 5 g

20000 ft 4 g at M=0.7 2.9 g, thrust limited

35000 ft 2 g at M=0.7 1.7 g, thrust limited

Flight Envelope +7.5 g, -3 g +8 g, -4 g

Ejection seats Zero - Zero Zero - 90 knots * Though not required by the RAF, it was proposed to offer a full strike version for export, having 5 pylons carrying up to 2268 kg (5000 lb) of external stores with two 30mm guns in under-fuselage pods. The maximum mass was estimated to be 7062 kg (15570 lb). The export market was estimated as 464 aircraft – the main competitor was judged to be the Franco-German “Alphajet”.


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1.5 Modified versions of the HS 1182. Soon after the issue of the brochure HSK 27, intensive negotiations were carried out by HSA and MoD, as a result of which, and with further investigation, changes were proposed, mainly with a view of reducing cost and eliminating any discrepancies. 1.5.1 The HS 1182 A The first of these changes was the ‘A’ version. It was slightly smaller all round than the basic aircraft, but still had full training and strike stores capability. The low speed tests on the half model at Hatfield had shown there was little to choose between the high and low wing positions from low speed considerations, but the 1/30th scale high speed model showed that, as the aerodynamicists had feared, a strong normal shock sat firmly in the channel between the intakes and upper wing surfaces and spread outwards as well. This gave an early drag rise and high drag at maximum Mach number. It also showed that the tailplane high position left much to be desired aerodynamically. On the model, filling in the channel to simulate a low intake gave much improved high speed results, but brought up fears of ingestion into the intake. The test results from this small high speed model brought about a major change to the layout of the aircraft. The intakes were moved down and forward, ahead of the wing root, and the intake side junction was carefully shaped so as to maintain wing isobar sweep to the fuselage side. There was a small increase in wing sweep to 26

o at the leading edge, and the rear fuselage was

‘banana’d’ down to put the tail into a more favourable position, assisted by anhedral on the tailplane. These changes ameliorated the high speed behaviour somewhat. The possible ingestion problem was addressed by analysing full scale test results of spray patterns on other aircraft and applying the results to the more forward low intake layout. This investigation gave promise that ingestion should not be a problem – this was later checked at full scale during tests on a very wet runway, with satisfactory results. The main undercarriage was moved to a position ahead of the main wing structural box, now an integral fuel tank. The move was necessary in order to accommodate the specified 30mm cannon on the centreline. The engine was moved back by 13 inches to balance. An example of the change of performance between the HS 1182 (HSK 27) and the 1182 A is shown in Table 3. 1.5.2 The HS 1182 V The possible need to drive down costs still further led to a reconsideration of the Viper 632-11 as a possible engine to power an even smaller version of the aircraft, in which the equipment fit was kept to a bare minimum, and there was no strike provision, only to weapon trainer standard. Wing area, span and fuselage length were all reduced and the basic mass also. Sized to meet the critical Sortie “E” of ASR 397, it met most of the performance aims, but with little margin for error, and, since there was no stretch left in the Viper, would be unattractive as a strike aircraft for export.


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Table 3. Performance of HS 1182 and 1182A against ASR 397 (iss.2)

Item ASR 397 HS 1182 HS 1182A Remarks

Take-off and Landing Performance

Not more than 4,000 ft MET MET

25 kn Crosswind MET MET

Approach 120-140 kn MET MET

Threshold 100-110 kn 112 kn MET

Speeds - Maximum - Max Continuous SL - Optimum Cruise

0.85M at 30,000ft 0.83 0.85

Not less than 420 kn 440 kn MET Limited by gust ride Not less than M = 0.6 0.7 MET

Climb with full fuel, time from brakes off to 30,000ft

Not more than 7min 8 min

7.2 min 60 % fuel

6.8 min

Ceiling Not less than 40,000 ft 45,000 ft 46,000 ft

Sustained manoeuvrability 4½ g at SL, 350 kn 3 g at 20,000 ft, M=0.7 2 g at 35,000 ft, M=0.7

5 g 5 g

2.9 g 3 g

1.7 g 2 g

1.5.3 The HS 1182 AT This version just replaced the Viper with the more powerful Adour, and not surprisingly showed a big step up in performance. But it had the same minimal standard of equipment as the ‘V’. 1.5.4 The HS 1182 AJ It seemed reasonable to use the extra performance available from the Adour to allow an increase in size and equipment standard to somewhere between the ‘A’ and the ‘AT’, and this resulted in the ‘AJ’ which closely represented the aircraft finally offered to the MoD (PE), and which won the design competition. Tables 4 and 5 show some details of all these variants.


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Table 4 Optional variants from HS 1182

Type HS 1182 A HS 1182 V HS 1182 AT HS 1182 AJ

Category Trainer / Strike Weapon Trainer Weapon Trainer Trainer / Export

Propulsion Adour Viper Adour Adour

Equipment standard

Full, with strike provision

Reduced Standard Reduced Standard Intermediate Standard

Length 11.64 m (38.2 ft) 10.12 m (33.2 ft) 10.12 m (33.2 ft) 10.33 m (33.9 ft)

Wing Span 9.77 m (32.04 ft) 9.0 m (29.51 ft) 9.0 m (29.51 ft) 9.4 m (30.83 ft)

Wing Area 18.02 m2 (194 ft2) 15.3m2 (164.7 ft2) 15.3m2 (164.7 ft2) 16.72 m2 (180 ft2)

Mass – Basic 3,465 kg (7,640 lb) 2,708 kg (5,970 lb) 2,971 kg (6,549 lb) 3,150 kg (6,944 lb)

Trainer, Max T.O. 4,795 kg (10,570 lb) 4,198 kg (9,255 lb) 4,209 kg (9,279 lb) 4,443 kg (9,794 lb)

Maximum T.O. 7,194 kg (15,860 lb) 4,835 kg (10,659 lb) 4,846 kg (10,863 lb) 6,898 kg (15,208 lb)

Bought out equipmt 842 kg (1,856 lb) 762 kg (1,680 lb) 764 kg (1,685 lb) 777 kg (1,714 lb)

BAMPR weight 1,949 kg (4,297 lb) 1,510 kg (3,330 lb) 1,524 kg (3,360 lb) 1,702 kg (3,752 lb)

BAMPR - explain

Table 5 Performance of optional variants.

Feature ASR 397 HS 1182 A HS 1182 V HS 1182 AT HS 1182 AJ

Max level speed at 30,000 ft, ISA, Mach no

0.85 0.85 0.82 0.85 0.85

Max continuous level speed, sea level, kn

420 440 434 500 500

Opt. range cruise speed at 30,000 ft, Mach no 0.6 + Met 0.6 0.65 0.7

Approach speed, 60 % fuel, knots 120-140 Met 125 127 130

Threshold speed, 60 % fuel, knots 100-110 Met 104 106 110

Time to 30,000 ft from brakes off, mins

7.0 6.8 8.5 5.9 7.0

Service ceiling, 60 % fuel, ISA, 13,400 m 40,000 ft

14,000 m 46,000 ft

14,300 m 47,000 ft

14,600 m 48,000 ft

14,300 m 47,000 ft

Sortie “E” duration hours 1.0 1.0 1.0 1.0 1.0

Sustained g, SL, ISA +15 o C, 60 % fuel, 350 kn 60% fuel, 20,000 ft, ISA, 0.7 M 60% fuel, 35,000 ft, ISA, 0.7 M

4.5 5.0 4.4 5.5 5.2

3.0 3.0 3.0 3.8 3.1

2.0 2.0 1.9 2.1 2.0

Take-off distance to 15 m height, still air, full fuel, S.L., ISA + 15 o C ft

4,000 3,150 3,500 2,500 2,800

Landing distance from 15 m height, still air, full fuel, S.L., ISA + 15 o C, wet runway ft

4,000 Met 3,870 3,900 4,000


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1.6 Possible Competitors A performance comparison was made between the HS 1182 and HS 1182A and various other training aircraft, and this is shown in the table below. Competing single and twin engine entries from Future Projects, BAC Warton (Project P.59) were not considered at this time, through want of information.

Aircraft Engine SL static thrust lb

T.O. weight lb

Minutes to 30,000 ft

Max level Mach no

HS 1182 Adour 172-06 5,000 10,220 7.2 0.85

HS 1182A Adour 172-06 5,000 10,374 6.8 0.85

Gnat T.1 Orpheus 101 4,400 9,217 6.2 0.95

Hunter Mk.7 Avon Mk.121 7,575 17,200 6.75 0.92

Macchi 326K Viper 600 4,000 9,680 8.0 0.82

Aero L – 39 Walter Titan 4,000 8,600 n/a 0.75

Saab 105 XT 2 x G.E. J-85 5,700 9,800 4.5 0.82

Alpha Jet 2 x T.M. Larzac-04 5,940 10,000 6.5 0.9 (?) Some comments on the MoD assessment are given in Appendix 2. 1.7 Draft Specification – 4th November 1968. By November 1968, the specification for the HS 1182 had become the following:

2 seats. Tandem (side-by-side also to be studied).

Simple and robust, capable of repeated high “g” applications.

World wide operation.

Provision for light weapons installation – not primary role. Performance

No external fuel or stores.

To be capable of operating from semi-prepared strips.

Take-off run 1600 ft, unstick speed 90kn. Concrete runway.

7 minutes to 30,000 ft.

Maximum speed 450 kn. at sea level.

Maximum level true Mach no 0.8+ at 30,000 ft.

Cruising speed 375 kn at sea level.

Maximum dive speed 500 kn / 0.85 Mach no.

Threshold speed 100 kn.

Sufficient fuel for 1 hour’s general handling sortie.


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1½ g sustained turn at 38,000 ft.

Landing distance from 50ft., wet runway, 3,600 ft. Handling.

To be flown by relatively inexperienced pilots.

Self limiting at high speed.

Cleared for aerobatics, stalling and prolonged spinning, recovery to be straight forward from these manoeuvres.

Capable of being flown solo from the front cockpit.

Stall to be clearly defined and with natural aerodynamic warning.

Mechanical warning acceptable only if a significant saving in cost. Structure.

Stressed to +8g, -3g design limits. Ultimate factor 1.5.

6000hr fatigue life on an agreed (severe) spectrum.

Anti-icing indicator, and engine anti-ice, but not intake, or airframe. Servicing, Reliability and Maintainability Standards.

To be laid down. This first issue was later modified as knowledge was gained, to emphasise the combat export version more strongly, to increase the maximum level Mach no. and to add -4g to the design envelope. 2.0 DEVELOPMENT OF THE HAWK T.Mk.1. 2.1 Contract negotiations The Company entered the competition to supply aircraft to fulfil the requirements of ASR 397 with a version of the HS 1182AJ. In effect this was approximately a 5 % smaller aircraft than the original HS 1182 on linear dimensions. For example, the wing area decreased from 200 sq. ft. to 180 sq. ft. However, items like the cockpit and engine were unchanged. The aircraft (later called the Hawk) was declared the winner of the competition in October 1971, and the contract and specification (281 D & P) were agreed on 12th March 1972. This was a fixed price contract for the design, development and manufacture of 176 aircraft (one to be retained for development) to an “Acceptance Standard” agreed between HSA and MoD (PE). All aircraft were to be built using production tooling – there were no prototypes. Ground and flight tests were to be monitored and verified by the Royal Aircraft Establishment (RAE) at Farnborough and by the Aircraft and Armament Experimental Establishment (A&AEE) at Boscombe Down – this was the usual process.


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The Acceptance Standard defined the airframe and equipment as “Contractor Furnished Equipment” (CFE) but the Adour engine was Government Furnished Equipment (GFE) and was the subject of a separate contract between the engine manufacturer and MoD (PE). This enabled HSA to negotiate directly with equipment suppliers, and also gave a firm basis for export negotiations. The export potential was recognised in the contract by a statement that the aircraft “should be capable of conversion to a close support role using one centreline pylon and two outer wing pylons, including the carriage of jettisonable external fuel tanks, heavy weapons and twin store carriers.” It was further stated that this capability should have a minimum influence on the aircraft as a trainer. Some very important amendments were made to the contract incorporating maintenance and reliability incentives, possibly for the first time in a MoD contract. Because these were novel, they are discussed in some detail below. 2.2 Incentives 2.2.1 Maintenance Target times were specified for some 95 maintenance actions, examples of which are shown below, together with the times achieved under controlled conditions. These were designed to ensure low life cycle costs. The targets were set from experiences on serving aircraft and analysis of actions in servicing. The demonstrations were carried out by normally trained servicing personnel.

MAINTENANCE EFFORT (man-minutes)

Action Specified

effort Demonstrated

effort Elapsed time

minutes

Pre- flight servicing 15 12.8 12.8

Turn round servicing 15 8.8 8.8

Post flight servicing 35 33.5 33.5

Re-arming (Gun pod & 2 pylons) 60 37.3 10.3

Engine change 500 369 103.8

Ejection seat replacement 240 42.3 35.8

UHF radio replacement 15 5.5 5.5

Battery replacement 15 7.0 7.0

To change or replace 15 equipment items (e.g. control column handle, trim actuator)

Average time 90

57.8 41


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All but two of the contracted baseline figures were achieved and HSA benefitted by the maximum payment allowed. 2.2.2 Reliability. The object of these incentives was to ensure a high reliability of operation when the aircraft entered RAF service. Targets were set for the maximum acceptable level of defects per flying hour on a range of airframe and equipment items. Two years after the Hawk entered service, defects occurring in 2,000 flying hours at the Advanced Flying School and 1,000 hours at the Tactical Weapons Unit were assessed. The achieved defect rate was far lower than the contractual base rate, resulting in a substantial payment to the company. Subsequent experience has shown that this good reliability has been maintained in service (300,000 hours by 1991) with defect rates and maintenance support at a significantly lower level than previously experienced on similar aircraft. This has provided a good cost saving, and confidence in maintaining an economic level of spares support. 2.3 Design and development. 2.3.1 Office organisation. On the receipt of the order, there began an intense phase of development on all aspects of the aircraft. Mr Gordon Hudson (ex-Folland Chief Stressman) was appointed Chief Designer and his Assistant was Mr K.G.Hodson, mentioned before. The Design Department was organised on the “matrix” principle, in which the line departments (Aerodynamics, Stress, Drawing Office, etc.) were responsible for all work within their discipline, but the various Chief Designers and Project Leaders could call for members of each department to be allocated to their project, full or part time as required. Not only did this make for better man-hour usage, but it also enabled a useful cross-fertilization of information between different projects to their mutual advantage. 2.3.2 Aerodynamic Development. Because this was the Author’s specialisation (appointed Head of Aerodynamics, Hawk in 1971 for ten years) this is dealt with in some detail. The Aerodynamics Department at Kingston relied heavily on analysis of wind tunnel testing to evolve the aircraft aerodynamically, both in the early days and throughout the flight testing. Significant configuration changes resulted from some of these investigations. Several wind tunnel models were built and tested:

(a) ½ scale, half model, high or low wing with flaps, undercarriage and tail (Figure 9), which was tested at low speed in the V/STOL wind tunnel at Hatfield. It was based on the early HS1182 layout, but was later modified to conform more closely to the final layout, when it became 9/16th scale to reflect the smaller wing area. It was extensively used to assess the high lift characteristics and, later, to investigate tailplane stalling – discussed later.


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(b) 1/30th scale high speed model, tested at Brough. This model gave the first indications of high speed problems due to interference between the high intake and the low wing, and the position of the tail. An ad hoc modification was made to simulate a low intake, which showed much improvement.

This resulted in the adoption of a low intake, moved forward to improve the chances of not ingesting spray from the nosewheel, which itself was moved back, to retract forwards into the nose. The rear fuselage was cambered downwards to bring the tail down, and the tailplane was given pronounced anhedral for the same purpose.

The model was modified to become 1/28.5th scale but was little used thereafter, being superseded by a later high speed model.

(d) 1/10th scale complete model (Figure 10), which was used for defining low speed aerodynamic characteristics in six degrees of freedom. This was the initial basis for the aerodynamic characteristics of the aircraft, following careful analysis of scale effects and compressibility effects. It was also used extensively to investigate high angle of attack data, for a preliminary study of spinning behaviour. It was later scaled to 1/9.5 to reflect the reduced wing area.

(e) Two-dimensional wing sections were tested at high speed in the ARA (Bedford) Two-Dimensional Tunnel (2DT) to form the basis of the wing design, and to fix the degree of sweep that was necessary to achieve the required drag rise Mach number.

(f) A new 1/6th scale low speed model was built for use in the 9ft x 7ft Woodford tunnel. This was the stand-by model used to investigate any aerodynamic problems that arose during the flight development of the aircraft, and to provide the basic body of aerodynamic information at low speed.

(g) A new 1/6th scale high speed model was tested at high speed in the 9 ft x 8 ft Transonic Wind Tunnel (TWT) at ARA (Bedford). This extended the results of the low speed model characteristics into the high subsonic and transonic regimes. Results obtained here showed that some help was needed to stabilise the shock pattern at around the drag rise Mach number, and to reduce the drag caused by boundary layer thickening. As a result of these findings, a programme of flight investigations was mounted, using and refining an array of wing vortex generators which successfully coped with the problem.

(h) The intake and internal duct of the aircraft was very carefully designed, and a 1/3rd scale model was built to ascertain its characteristics. It was tested at Kingston under static conditions, and in the ARA Transonic Wind Tunnel with exterior flow. In addition, a full scale intake with an engine was run statically by Rolls-Royce.

(i) Two small (1/18th scale) lightweight models were built and tested in the spin tunnel at IMFL Lille. These were simple models; the first had pre-set controls, but the second had a two-position setting (pro-spin and recovery) which could be activated by radio control. This latter model was also tested with a variety of store configurations. The results indicated that for consistent spinning on the models, some form of strake was needed on the nose, but this was never tried on the aircraft, and the spinning behaviour in trials (over 800 spins) was judged satisfactory for training.


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(j) Some years later (probably in 1979 or 1980), a 0.315 scale model was built for testing in the newly commissioned RAE (Farnborough) 5 metre, high Reynolds number, wind tunnel. With the tunnel pressurised to about three atmospheres, stalling flight Reynolds numbers could be achieved, as well as the correct Mach number. This was a joint programme with the RAE who built the fuselage, with the flying surfaces built at Brough. There were separate testing programmes by HSA (mainly for high lift work on the USN T45) and the RAE for testing with pylons and stores. This model was also used by HSA in the 13ft x 9ft low speed wind tunnel at Weybridge, again for T45 work. A novel test used gauze screens upstream to increase turbulence, in order to simulate high Reynolds number, with some success.

Mr Barry Pegram, then Section Leader in the Fluid Dynamics Group in the Aerodynamics Department, was very closely involved in the wind tunnel activities throughout the aerodynamic development of the Hawk, and in the associated flight test analysis. Much of the above information is due to him. He also mentioned another model, often forgotten. When the Swiss Air Force were considering placing an order for Hawks, he visited the wind tunnel at F&W (Emmen) to see if some offset work could be placed there. A small high speed wind tunnel model was built and tested producing data for wingtip-mounted Sidewinder missiles, which information was used, it is believed, by BAe Brough, who took over the Hawk project from about 1986 onwards. Many of these models were used alongside flight development testing, often to investigate some quirk of behaviour found during a flight test. Very often a test in the wind tunnel indicated a possible solution, which was subsequently tried out and developed by further flying. More details of the model testing are given in Appendix 1 (a). The following sections give an account of a few items in flight development, mainly of aerodynamic interest. 2.3.3 The engine air intake Some notes on the principles behind the very good air intake of the Hawk may be of interest. The inlet duct highlight area had a contraction ratio of 1.3 to the throat, which was sized to give a mean Mach number of a modest 0.5 at maximum air mass flow at take-off rating, sea level ISA, in order to have something in hand for when the engine might be uprated in the future. All the bends were at constant flow area, and the expansion of area was done along straight sections of the duct. The splitter went close to the engine face, where the local mean Mach number was about 0.4 under the same conditions. The measured flow distortion at the engine face was small and the intake performance was very good, 97.5 % pressure recovery under static conditions at take-off rating. It was probably in large part due to the excellent design of the intake by Kit Milford that the Adour engine had no handling restrictions in normal flight, and only in heavily stalled or spinning flight was some care necessary.


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The relight envelope was wide and relights were made without difficulty. 2.3.4 Flight Development. Duncan Simpson, Chief Test Pilot, flew the first flight from Dunsfold on 21st August 1974 (Figure 13) and this was followed by a series of flights by other pilots for familiarisation and expansion of the flight envelope. For example, the third flight was carried out by Andy Jones, the nominated Project Pilot.

Figure 13 Hawk T.Mk.1 first aircraft (XX 154), first flight 21st August 1974

It is not possible within the scope of this document to give anything like a full account of the flight testing that was undertaken by four development aircraft (XX 154, XX 156, XX157, XX158) and the company demonstrator ZA 101, G-HAWK, (Figure 14). Their initial flight dates and number of flights made were as follows: XX 154 21.08.74 489 flights 405hr 30min up to 18.1.82

XX 156 19.05.75 1207 flights 1,184hr 10min 29.04.88

XX 157 22.04.75 232 flights 201hr 59min 21.02.76

XX 158 01.07.75 380 flights 313hr 57min 18.07.78

ZA 101 17.05.76 N/A flights 1,754hr 20min 14.11.88 Other aircraft used from time to time were:

XX 159 17.06.75

XX 160 19.11.75


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The crews for these first flights were:

XX 154 Duncan Simpson (solo)

XX 156 Jim Hawkins & Mike Snelling

XX 157 Andy Jones & Jim Hawkins

XX 158 Andy Jones (solo)

ZA 101 Duncan Simpson (solo)

XX 159 Mike Snelling & Chris Roberts (?)

XX 160 Jim Hawkins & David Young See also Appendix 1(b)

Figure 14 Hawk Test Aircraft (XX 156, XX 157, XX 158) at Dunsfold Airfield Initial comments were pretty favourable on both handling and performance within the limited flight envelope then available. In particular, the cockpit was described as outstanding. As the flight envelope was pushed out towards its specified limits, some aerodynamic deficiencies started to become apparent, and required correction, though it is fair to say that none were show-stoppers requiring serious re-design. The next section discusses some of the less desirable behaviour items found during the flying, and how they were put right 2.3.5 Some particular flight test events. (a) Stall behaviour The stall, as first experienced, occurred at a good low speed, but with very little buffet warning. One or other wing dropped suddenly and uncontrollably, though the aircraft did not depart into


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a spin and recovery was normal, rolling out to controlled flight. But clearly this was not acceptable. Flow visualisation revealed that a sudden ‘leading edge’ type of stall was occurring, originating at about mid-semispan. This was perhaps partly due to an aerodynamic concession, in that for production simplicity (and hence lower cost), the trailing edge flaps had been made with constant chord and section. But they were fitted to a fairly sharply tapered wing (taper ratio 0.34. tip to root) and so the ratio of flap chord to wing chord was at its highest at the outboard end of the flap and required too much of the flow at the local leading edge. Guided by work on the half model, the flap vane on the outer part of the flap was removed to detune it somewhat, at the cost of some maximum lift. Although the initial separation of the flow still occurred at the same point, the flow breakdown at the stall was kept from rapidly spreading towards the outer wing by the judicious positioning of a large fence. Buffet warning was obtained, at the cost of a little more maximum lift by putting triangular section “breaker strips” on the leading edges, inboard to give warning and outboard to give repeatability. As is related later, the outboard end of the flap vane was removed for another reason, and together with the devices above now gave acceptable behaviour, but lost about 5 knots of stalling speed. However, there was enough of a margin in maximum lift coefficient to meet the field performance requirements for the RAF. Clearly there was scope for much more fine tuning and investigation of more refined stall fixes, but there was a tight deadline to meet for the RAF, and this work was left to be done on the later developments. (b) Howling and the Phantom Dive. During one of the early stalling flights, before the flap vane had been cut back, two curious phenomena had been noticed. The first of these was a report from the pilot that when the flap was travelling from ‘up’ to ‘mid’ there was an intermittent ‘howl’. Now the flap vanes were fabricated from glass-reinforced plastic, and it quickly became clear that at an intermediate position the local internal airflow was causing them to vibrate. This was easily cured by putting in more stiffeners between the vane and the flap. The second of these was more serious and demanded immediate attention. It was first discovered when recovering from a stall with full flap and undercarriage up. It was found that at forward centre of gravity in that configuration, rapid fore-and-aft movement of the control column could induce an uncontrollable nose down pitch, with the nose down attitude and speed increasing quite rapidly. Recovery was straightforward, either by retracting the flap a few degrees, or by extending the undercarriage, but this was not acceptable as an operation, even though the configuration was unlikely to be used normally. It was dubbed the “Phantom Dive”. This term had been coined after an initially unexplainable series of fatal approach accidents on the Gloster Meteor, because in those events there was a sudden loss of control under conditions which were normal and correct for final approach (“It came like a phantom, from nowhere”), and the aircraft dived into the ground from low altitude. After investigation it was found that


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on the Meteor, selection of airbrakes when the flaps and wheels were down, a seemingly logical operation, gave rise to an interference which caused the tail unit to become ineffective. It is believed that a tailplane stall had also been experienced on the F-4 “Phantom” but the expression did not derive from that. It was shown with the half model of the Hawk at Hatfield that high local downwash at the tail, coupled with the very large nose down pitching moment induced by the flap, was causing the tailplane to stall on its lower surface, so that it could no longer provide adequate balancing power. It needed more lift, extended to higher angles of attack. In the case of the F-4K Phantom aircraft this was achieved by installing a fixed leading edge slot to the tailplane, harking back to demonstrations of such devices by Handley Page on wings in the early Twenties! A fixed slot with its associated drag was not an option on the Hawk although a cambered tailplane was tried on the model with some success. Removal of the outboard vane of the flap reduced the flap pitching moment to such a value that the standard tailplane could cope, so this was the quick solution for the RAF. However, for the US Navy VTX project, (and for later combat versions of the Hawk) the maximum possible lift was required, so that at least the outer flap vane had to be replaced. The dive phenomenon had to be fixed. An example of the cross-fertilization of knowledge due to the matrix working of the design department now occurred. Barry Pegram, then Section leader of the Fluid Dynamics section of the Aerodynamics Department, had been working on the adoption of leading edge root extensions (LERX) for the “Harrier” wing, and this work had shown that these devices extended the lift of the wing to higher angles of attack by virtue of the non-linear lift developed by the vortex flow they created. He proposed that these should be added to the tailplane of the Hawk model, but this would have had a serious effect on the tailplane hinge moments. The author, who was working with Barry in the V/STOL tunnel at Hatfield, suggested that the ‘tailplane canard vane’ (TCV), as it was called, could be fixed to the fuselage at such a position that it was lined up with the flow at normal conditions, but with its trailing edge adjacent to, with a small clearance, the leading edge of the tailplane at its maximum nose down position. Experimenting showed that these vanes could be made quite small and they gave a complete cure to the problem in the wind tunnel, with very little drag in the normal flight regime. To prove the concept in flight, some temporary vanes were manufactured which could rapidly be fitted to one of the test aircraft which had flaps with the full vane. First, the aircraft was flown without the TCV to establish the conditions under which the ‘Phantom Dive’ occurred on that particular aircraft. On a later flight, the TCVs were fitted and the aircraft flown again to the identical conditions as before. Despite every attempt by the pilot to instigate the phenomenon, it did not occur – the vanes were a complete success, even though they looked inconspicuously small for such a large effect. The aerodynamicists had a surprisingly difficult job to ‘sell’ the idea. Before the use of the TCV was sanctioned, tests were demanded with the high speed model at ARA and with the spin models at Lille. They were shown to have negligible effects under all normal conditions. Later, their effectiveness was again demonstrated using the 0.315 scale model of the Hawk in


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the RAE (Farnborough) 5 metre, high Reynolds number, wind tunnel. They became standard for the T45 aircraft, and for the 100/200 series of the Hawk. (c) Directional wander. Another feature that was adversely commented on in the early days, as experience grew with the aircraft, was a premature low level buffet, generally felt on the rudder, and a tendency for the aircraft to “wander” from side to side over a degree or two of yaw at attack speeds, making weapon aiming an unduly difficult task. This could be controlled by very firm fixing of the (manual) rudder by the pilot, but this was undesirable. The reason for both the buffet and the wander was revealed by flow visualisation around the area aft of the fin, where the fuselage top profile was quite rounded. It was shown that periodic shedding of the boundary layer occurred in the area, and this induced the buffet and the motion. Fred Sutton, Head of the Flight Test Department at that time, suggested that the upper line behind the fin be raised, and the edges be sharpened so as to fix the separated boundary layer. The modified shape was informally known as “Fred’s Back End”; it proved successful in curing the buffet, and was adopted. The directional wander was found to be due, at least partially, to the wake shed by the two air conditioning inlets behind the cockpit canopy. Nothing could be done about these, but fitting ventral fins either side of the airbrake, coupled with the adoption of a stiff centralising spring in the rudder control circuit, was judged to give acceptable behaviour. The aerodynamicists would have liked to have an irreversible, fully powered rudder, but this was deemed to be too expensive. It was also demonstrated that there was a directional trim change with power. This was eventually traced to leakage flow from the ‘bacon-slicer’ seals between the rear fuselage and the root of the tailplane. The problem was much ameliorated when the sealing was improved. Early on, the ailerons were found to be rather sensitive and, together with the directional wander, made weapon aiming a very difficult task. A non-linear gearing was introduced to reduce aileron angle per inch of stick movement in the middle range. (d) High speed flying. This seemed to be pretty good up to the drag rise Mach number, but this was a little earlier than was forecast, and there was some roll uncertainty between M = 0.8 and 0.9. Tests with the ARA high speed model showed that the local wing shock system was fairly strong and might be causing boundary layer breakaway on one wing or the other, which would explain the roll uncertainty and possibly the earlier than expected drag rise. The attack on this in flight started with an array of numerous large, rectangular, vortex generators situated on the upper wing surface just ahead of where the shock would be sitting. This was successful, but faced with the threat by the Chief Engineer that the author’s salary would be reduced by five pounds per week for every vortex generator on the wing, these were very quickly reduced in size and number until the minimum array consistent with acceptable results was reached. This consisted of eight small vortex generators across the wing, four either side of the wing fence. In addition, the vortex generators increased the aileron effectiveness at M = 0.9 and above.


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Thus equipped, the aircraft was demonstrated in powered dives during which it was found to be self-limiting to a Mach number of about 1.17. (e) Rolling pull-out tests. These are often quite hazardous for an aircraft, taking it to the limits of such cases as fin strength. On the Hawk, the critical parts of the structure were strain-gauged and calibrated, so that the stresses imposed by these flight manoeuvres could be monitored flight by flight, progress being made by incremental increases in applied ‘g’ and speed. The test programme showed the expected increase of stress with increase of applied ‘g’ but extrapolation to higher ‘g’ began to look limiting. Fortunately the stresses became non-linear and levelled out a bit, increasing at a slower rate. The required ‘g’ and speed limits were reached with some margin of strength in hand. (f) Engine handling clearance. A number of flights were devoted to investigating the handling and relight envelopes for the Adour engine. No great difficulties were encountered and a very useful relight envelope was cleared. There were no handling restrictions on engine handling over the whole flight envelope, with the exception of spinning and flight at very high angles of attack. But it was noted that the engine acceleration was on the slow side compared with straight jets. 2.3.6 Flying by A&AEE, Boscombe Down, 1975 – 1976. Testing by Boscombe Down pilots to ASR 397 requirements was carried out during the general programme, once the configuration was more or less fixed. Their results were generally very favourable. XX 159, XX 157, XX 160 and XX161 all took some part. A ‘letter report’ was issued on 30.07.76, which stated “The aircraft…ideally suited for its intended role as a flying trainer.” The performance requirements of ASR 397 were all met or exceeded. There were adverse comments on the buffet at M=0.7 and above, and the directional wander, and there was an unacceptable trim change with airbrake extension above 450 knots. This was before the actions discussed earlier, and the airbrake problem was ameliorated by altering its shape, and partly balancing the tailplane operating rods, which were affected by deceleration. A modification to put in an automatic link to apply a small nose-up tailplane movement was engineered but not implemented. The assessing crew agreed that the ‘Phantom Dive’ problem had been completely cured by cutting back the outer flap vane, and the other wing dressings. Figures 15 and 16 show the T.Mk.1 upper and lower wing surfaces.

Figure 15 Hawk T.Mk.1 Upper surface view Leading edge breaker strips and fence just visible


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Figure 16 Hawk T.Mk.1 Lower surface view Leading edge breaker strips and fence visible Faired flap hinges and aileron rod fairings towards trailing edge of wing

2.3.7 Subsequent events. Initial C.A. Release was achieved on 18.10.76 and a second issue on 14.07.77. Deliveries began with production aircraft XX 162 and XX 163 flown to RAF Valley (No.4 F.T.S) on 4th November 1976, 27 months after first flight and 4½ years after contract agreement. In September 1976, nine Hawks appeared in formation at the SBAC air show at Farnborough, the first large formation of Hawks seen in public (Figure 17).

Figure 17 Nine Hawks on the way to the SBAC Airshow, Farnborough 1978


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At about the same time, in a letter to HSA from the Ministry of Defence, the following statement was made:

“HSA are to be congratulated that the Hawk is on time and on cost, and has met and/or exceeded the performance specified.”

In service with the RAF, the Hawk was mainly used for Advanced Flying Training at RAF Valley (Figure 16) and for Tactical Weapons Training at the Tactical Weapons Unit at RAF Brawdy and Chivenor (Figure 18). The Red Arrows RAF Display team took on the Hawk T.Mk.1A (Figure 19), disposing of their Gnat Trainer aircraft. They used a slightly modified version of the Adour, which had a faster response time than the standard engine. Later (January 1983), some 88 Hawks were rewired so as to be quickly convertible to a second line defensive fighter role in an emergency, carrying a centreline gun pod and Sidewinder air-air missiles on the pylons (Figure 19).

Figure 18 Hawk Weapon Trainer Figure 19 Foreground – Red Arrows with Matra rocket projectile packs Hawk with centreline smoke fuel tank Background – up to 88 Hawks, including

Red Arrows, could carry Sidewinders 2.4 Structural Testing 2.4.1 Static strength tests Static strength tests (Figure 20) were made at Kingston on various components, and on the complete airframe, for a number of symmetric and asymmetric cases. A buckle occurred at frame 29 at 67% of fully factored load (FFL), and was easily repaired and strengthened to reach 100% FFL in that particular case and 120% FFL in a related case (fuselage asymmetric loading). The fuselage fuel tank area went to 109% FFL. With the wing, some localised failures occurred on ribs 8 to 10, at the leading edge and the first integral stiffener aft of the front spar, at 57% span. Again these were easily corrected. Eventually the wing failed under symmetric loading at 133 % FFL. Clearly this was a very strong airframe, and capable of going to much higher weights in the developments which carried attack stores.


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Indeed, for the export variants, total store loads of up to 3000 kg were demonstrated, typically five 500 kg bombs, one on each pylon.

Figure 20 Static strength airframe under test Ground test Services (Kingston)

2.4.2 Fatigue strength tests. A comprehensive series of accelerated fatigue testing was instituted, and proceeded for some time, keeping ahead of aircraft flying hours. The fatigue spectrum to which the airframe was subjected was very severe (see Figure 21) and there was every indication that the requirements of the specification would be easily met.


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Figure 21 Comparison of Hawk design fatigue spectrum

with Jet Provost and Strikemaster 2.4.3 Noise testing. Though not strictly a structural testing item, it had been found in flight tests that the cockpit environment was acoustically noisy, especially above 450 knots, so much so that with ordinary helmets intercommunication was difficult above 500 knots. There was much discussion with RAE Farnborough about this. Measurements in flight showed that the main source of excitation of the cockpit side panels was the area near the nose of the external boundary layer diverter between the inner side of the intake and the fuselage side. It was probable there was a horseshoe vortex around the rounded nose of the diverter. Nothing much could be done aerodynamically, but two solutions were discussed. The first of these involved lining the whole cockpit wall area with adhesive acoustic tiles shaped to fit between the stringers and frames. These were very heavy and required an inordinate number of man-hours to fit, but some sets were ordered and some tried out with some success. But they added a great deal of unwanted weight. The other solution was to accelerate the introduction of improved “bone domes” having better sealing and a much better acoustic performance. This was in hand anyway for other aircraft, and was much cheaper in the long run, so this was the procedure adopted. The few sets of acoustic tiles that had been delivered were scrapped, and a lot of the cars around Dunsfold suddenly became a lot quieter ! Dr John Green, who was Head of Noise Division at RAE from 1975 to 78 and chairman of the Cockpit Acoustics Group, has added the following comment.


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“When the problem of the Hawk cockpit noise surfaced, the Cockpit Acoustics Group had immediate access to the Boscombe Down noise measurements. We thought at once of flow separation ahead of the intake diverter and backed this up with some quick measurements in flight. The solution in this case was not to modify the diverter but to stick lino tiles on the skin underneath the vortex created by the separation ahead of the diverter. I remember that Ken Heron was adamant that the purpose of the tiles was solely to add mass to the skin in order to reduce transmission, and that only the skin under the vortex needed treating. Until I read Harry’s paper, I had nursed the belief that every Hawk was flying around with our lino tiles by the pilot’s knee to keep the noise down. However, while all this was going on I was chairing a sub-committee of the Cockpit Acoustics Group that had the task of drafting a specification for allowable noise levels in the cockpits of military aircraft. We devised a procedure that started with the noise spectrum in the cockpit, subtracted from it the attenuation spectrum of the helmet, added in the spectrum of voice communication at a signal to noise ratio of 10dB and tested the resulting spectrum against the risk of permanent noise-induced threshold shift (PNITS) over an operational service life. This procedure went into the design requirements document Av.P. 970. The Mk V helmet was new at the time and, when set against the requirements in our newly devised specification, it seemed likely to be the solution of the cockpit noise problem. I was unaware until now that this process had resulted in the discarding of tiles in the Hawk cockpit.” 2.5 A brief description of the Hawk T.Mk.1. 2.5.1 Crew Station As can be seen in the illustrations, the Hawk has a two seat tandem cockpit with stepped rear seat, giving outstanding vision to both pilots. The rear position, usually occupied by the instructor, gives him a view of the runway ahead almost to touchdown and also permitted the use of a weapon sight for training purposes. The one-piece curved stretched acrylic windscreen gives good rain clearance and absorbed the impact of a 2lb bird at 450 knots (later 528 knots with a modified windscreen). The main one-piece canopy is also made of stretched acrylic material and is hinged sideways for entry. The canopy is fitted with miniature detonating cord (MDC) which is used to shatter it, prior to ejection using Martin Baker Type 10B Zero-Zero rocket ejection seats. These take only 0.7 second to launch the crew in their seats. Command ejection could be selected from the rear cockpit.

Great care was taken with the Figure 22 Hawk Front Cockpit Instruments instrument layout. (Figure 22) (key on following page)


37

Key to instruments and controls, Figure 22

The double air conditioning pack sat just behind the canopy and delivers pressurisation of 4psi differential using engine bleed air. Oxygen supply comes from two charged bottles. 2.5.2 Fuselage. The fuselage is of conventional aluminium alloy stringer – frame construction. Ahead of the cockpit, the nosewheel retracts forwards into the nose compartment, which provides storage for some of the equipment. A MicroTurbo 047 Mk.2 gas turbine driving an air compressor supplies air to start the main engine, which is mounted on two forward and one rear brackets. The engine can be taken out of the airframe by using mini - hoists to lower it on to a trolley beneath the aircraft. A bag-type fuselage fuel tank of 191 imperial gallon capacity is positioned in the fuselage over the wing and is pressurised by bleed air. Much of the fuselage surface is covered by removable panels, designed for rapid access to equipment items for servicing or replacement (Figure 23). This is a significant factor in reducing maintenance man hours per flying hour. The lower part of Figure 23 re-iterates the design philosophies employed on the Hawk.


38

Figure 23 Hawk airframe access areas and design features A gun pod housing a 30mm Aden Mk.4 cannon and 120 rounds of ammunition is carried on the centre store point. In the Red Arrow aircraft, this is replaced by a tank containing the oil used to create smoke by injecting it into the engine efflux from pipes just above the exhaust. 2.5.3 Wing The one-piece wing, made from aluminium alloy, has continuous structure from tip to tip, and is attached to the fuselage by six bolts. The main box has machined skins with two main spars and integral ribs and stringers. The whole box structure is used as an internal fuel tank, with a capacity of 184 imperial gallons. An auxiliary spar at the front of the wing forms the front of the compartment holding the retracted main wheels.


39

The ailerons are also from aluminium alloy, but are filled with honeycomb. Most of the many inspection hatches are fitted on the top (compression) surface of the wing, with ‘form-in-place’ seals for good surface finish. For the RAF aircraft, only inner wing pylons are provided, capable of mounting weapon trainer stores up to 1,500 lb weight, but export versions of the aircraft carry up to five wing and centreline pylons, the inner ones being capable of carrying external fuel tanks with capacities of 100, 130 and 190 imperial gallons each. The maximum weight of external stores in this case could be up to 6,800 lb (which has been demonstrated). 2.5.4 Tail unit Again this is of conventional aluminium alloy construction. The rear tailcone incorporates a housing for a 2.64 m (8ft 8in) braking parachute as an option. 2.5.5 Systems. The hydraulic system operates at 3,000 p.s.i. and powers flying controls, flaps, undercarriage, airbrake and anti-skid wheel brakes. There is a “pop-out” ram air turbine for emergency use to power the flying controls. The flying controls, with the exception of the (manual) rudder are powered by duplicate irreversible hydraulic jacks, the valves of which are linked to the cockpit controls by push-pull rods and links. There is a bob-weight in the nose to control the longitudinal stick force per “G” and feel is by springs and non-linear gearings for tailplane and ailerons. An engine-driven DC generator plus two stand-by batteries provides the electric power, with inverters for AC equipment. Gaseous oxygen is carried in two bottles. Avionics normally include VHF, UHF TACAN, glideslope receivers and IFF/SSR. 2.5.6 Powerplant. The powerplant is the Adour Mk. 151 - 01 in the RAF. The Red Arrows use the Mk.151 – 02 which has a modified fuel system device to shorten the engine response times. An illustration of the engine, and more details of its performance, and of its variants, are given later. 2.5.7 Undercarriage. The main units are mounted outboard on the wing and retract inwards and forwards into the front of the wing root and the adjacent fuselage. They have single wheels and a levered suspension, trailing link system. The castoring (non-steering) nosewheel also has a similar suspension layout with a single wheel, usually fitted with a ‘chine’ tyre to assist surface water dispersal.


40

2.5.8 Leading dimensions of the Hawk T.Mk.1.

Note These dimensions can be compared with those for the HS 1182 (HSK 27) project listed on page 14.

2.5.9 Weights. Basic Mass (Weight) 3,450 kg (7,606 lb) 2 crew 172 kg (380 lb) Fuselage fuel 841 litres (185 I.G.) at s.g = 0.79 664 kg (1,462 lb) Wing fuel 864 litres (190 I.G.) at s.g.= 0.79 682 kg (1,504 lb) Role equipment 249 kg (549 lb) Max Mass (Trainer) 5,000 kg (11,000 lb) 2.5.10 Performance. Symmetric flight limits, cleared for the R.A.F for the T.Mk.1 are shown in Figure 24.

Wing

Span 9.39 m (30ft 10in) Aspect Ratio 5.284

Reference area 16.69 m2 (179.6 ft2) Taper Ratio 0.34

Sweep at leading edge 26 o Sweep at ¼ chord 23½ o

Section thickness/chord ratio

At root 10.9 % At extended tip 9 %

Dihedral angle 2 o Wing setting 1 o

Pylon centre lines

inboard 2.399 m (7ft 10in) outboard 3.422 m (11ft 3in)

Fuselage overall length 11.459 m (37ft 7in)

Tailplane Tail Arm 4.3 m (14ft 1in)

Span 4.39 m (14ft 5in) Aspect Ratio 4.45

Reference area 4.328 m2 ( 46.6 ft2) Taper Ratio 0.33

Sweep at leading edge 34.6 o Sweep at ¼ chord 30 o

Section thickness/chord ratio

8.5 % Dihedral angle - 10 o (anhedral)

Fin and Rudder

Height above fuselage 1.772 m (5ft 10in) Fin arm 3.492 m (11ft 5in)

Reference area 2.508 m2 (27 ft2) Section thickness/

chord ratio 8 % to 9 %

Sweep at leading edge 45 o Sweep at ¼ chord 39 o


41

Rolling pull-outs 720 deg. roll 1 “g” 200 to 500 kn or 0.81M 360 deg. roll 0 to 3 “g” 200 to 500 kn or 0.81M 180 deg. roll -1 to 2.5 “g” 200 to 425 kn -2 to 5.3 “g” 425 to 500 kn or 0.81M Cleared for up to 4 turn erect spins (but up to 13 turns demonstrated). 1st A&AEE assessment against ASR 397 (XX 154, 14th to 22nd April 1975) ASR 397 Achieved

Climb from brakes off to 30000 ft. Not greater than 7 min. MET Max. Level Mach no. at 30000ft. 0.85 (0.81 Acce. Std.) 0.87 Level speed at 2000 ft. 420 kn at S.L.(475 kn Acce.Std.) 521 kn. Thrust boundary at 2000 ft. 350 kn. 4 “g” 6 “g” at 20000 ft M=0.7 3 “g” 3 “g” at 35000 ft. M=0.7 2 “g” 2.2 “g” The most critical, sortie “E”, with reserves 1 hour MET Take-off and landing to/from 50 ft (wet runway) 4000 ft. MET

Figure 24 Flight Limits for the Hawk T.Mk.1 - R.A.F. service.


42

2.5.11 Aircraft drag (clean) An analysis using brochure engine performance and measured speeds at various weights, altitudes and engine power has been made, and the following partial drag polars have been derived by the author, and presented for interest:

CD = 0.0204 + 0.1079 CL2 at M = 0.7 up to CL = 0.57

CD = 0.0204 + 0.1089 CL2 at M = 0.8 up to CL = 0.28

or CD = 0.0169 + 0.1528 CL2 at M = 0.8 above CL = 0.28

N.B. These derived drag figures may not be the actual drag, but used with brochure engine performance tables they should reproduce the flight-measured level speed performance. If the actual engine performance were lower than the brochure, then the derived drags would also be lower. 2.6 Flying the Hawk. 2.6.1 Service Acceptance of the Hawk. One cannot do better than quote from a statement by Brian Hoskins, well-known former Leader of the Red Arrows. He converted to the Hawk in the summer of 1979, and the team started flying the aircraft intensively in October of that year. They started displays in the aircraft the following April. In 1981he wrote: “The Hawk is undoubtedly a more advanced plane than the Gnat, and it’s a very comfortable one to fly. Its major advantages are that it is extremely reliable, and both carries more fuel and has a far more efficient engine than the Gnat. As a result, we can get much greater flexibility in diversion: we can finish one display and go on much further than we could before. Last year, we did more than 120 displays, and the aircraft performed very well indeed. I cannot envisage any limit on the time we shall use the Hawk. I would have thought it will be in service for very many years, and will be flown by the Red Arrows for a long time to come. The displays we do with the Hawk are essentially the same as we did with the Gnat. But we do need rather more anticipation than we needed with the Gnat; the Hawk has forced us to change our technique a bit, especially with the throttle. In particular, you need to use the airbrakes against power much more than in the Gnat. Another feature of the Hawk is that it is an excellent trainer. It is supersonic, you can fly it on long sorties, you can spin the aeroplane and, of course, it is very easy to handle in aerobatic formation. A real advantage is that, as well as serving as an advanced trainer, you can use it for weapon training. It’s a marvellous turning aeroplane; its wing is very strong and produces plenty of lift; and when you get into a Hawk the good thing is that it feels as if the entire plane really has been designed to help you do your job. Everything is as it should be to make it easy


43

for you to fly. In fact, I would think that the Hawk’s only fault as a trainer is that it may be a little too easy to fly.” (Extracted from “Royal Air Force – Aircraft in service since 1918” Published 1981 by Hamlyn Publising Group, Astronaut House, Feltham, Middlesex, U.K. ISBN 0 600 34933 0) 2.6.2 A British View. There were many column-inches of copy from aviation journalists writing in a number of periodicals, when the Hawk data were released. One of the most informative articles, “Poised to Strike” written by John Fricker, was contained in “Air International” issue September 1978 (A Fine Scroll publication, De Worde House, 283 Lonsdale Road, London SW13 9QW),. 2.6.3 An American View. Maj. John P. Kelly, USAF was on an exchange posting to RAF Brampton when the opportunity came to fly the Hawk. He was a Senior Pilot and a flight instructor in the USAF’s Air Training Command. He had completed a tour flying RF-4C aircraft with the Tactical Reconnaissance Squadron and had 4600 hours to his credit. He wrote the article “Hawker Siddeley’s Hustling Hawk” which was included in the June 1977 issue of AIR FORCE Magazine, published by the Air Force Association, Washington D.C. He was fulsome in his praise of the aircraft. 3. HAWK VARIANTS 3.1 Preamble From the beginning it was always recognised that an RAF order alone would not cover the costs of the project to HSA. It was clear that alternative versions would have to be developed from the initial version, and these must be offered for sale to other Air Forces. In particular Scandinavia, with the exception of Sweden who had their own vigorous industry, was a possible source of orders, and also the Middle East, together with the emerging nations of Africa. The Far East was also a possibility, with Malaysia and Indonesia as front runners. It was further recognised that many of these countries wanted a dual-purpose role for the aircraft, a trainer without stores which could quickly be converted to a light strike role, carrying external stores on multiple pylons. Thus the schemes were designed to have a four- or five-pylon capacity, and the structure had to cater for these in terms of maximum weight and equipment. The MoD Specification for the Hawk explicitly mentioned this, but warned that any impact on the T.Mk.1 should minimal. The next sections describe the variants that evolved from the original design, and briefly describe their features. The VTXTS project of the U.S. Navy, which ultimately led to the largest single order for any Hawk variant, is mentioned only in passing. The author was involved with it only at the


44

beginning, and feels it is best left for those more deeply involved to tell its part in the Hawk story. However, there is a more detailed coverage in Part 2. The Hawk continues to be developed, with new orders received. This Section covers a period up to 1995, the 21st anniversary of the first flight of Hawk XX 154 – still flying ten years later, it is believed. 3.2 The Mk.50 series The Mk.50 series was closely related to the T.MK.1, with the same Adour engine, now the export version, the Mk.851, but it was fitted with “wet” inboard pylons, capable of carrying 130 imperial gallon external fuel tanks, which could be jettisoned if required, and outboard pylons. A central pylon was an option if the gun pod was not used. All the pylons were stressed to take stores of up to 500 kg nominal, and smaller stores could be carried on twin store carriers on the wing pylons. Each pylon was equipped with Ejector Release Units (ERU), to ensure clean separation from the aircraft. A fairly comprehensive weapons management system was offered, and enhanced navigation and attack avionics, including new instruments (for example an angle of attack indicator) and improved communications. The cockpit layout was improved with additional instruments. Dimensions were much the same as the basic aircraft, but the nose equipment bay was slightly larger. The flying surfaces were all the same size, but because of the higher weights, an effort was made to improve the high lift capability. To help reduce the landing run, particularly when heavy stores were returned to base, an 8ft 8in diameter brake parachute was offered, contained in a modified tailcone. Upgraded wheels and tyres were fitted.

The T.Mk.1 had been given a ‘quick fix’ to the stall behaviour in order not to delay delivery. This consisted of an outboard fence and ‘breaker strips’ inboard and outboard. These cost perhaps 5 knots of stalling speed, but there was sufficient in hand for the RAF field performance requirements to be met. Though the initial Series 51 deliveries retained the standard wing dressing, a programme of flight tests to improve maximum lift commensurate with acceptable behaviour was instituted on the Company demonstrator G-HAWK (ZA 101). Eventually an arrangement was developed which put a small breaker strip just inboard of the wing fences, and three small subsidiary fences were located on the wing surfaces inboard, at approximately equal spacing (see Figure 25). This became standard for the Mk.60 series, dealt with in the next section. Figure 25 ZA 101 as 60 series development aircraft Note wing dressing of three small fences


45

G-HAWK also had the task of flying, dropping and clearing a multitude of different stores including 100 and 130 imperial gallon tanks. These were originally as flown on the Hunter, but it was found that the tail fins were buffeted by the wake, and the whole tank tail cone was removed, the resulting small base area adding very little drag. Empty fuel tanks are notoriously prone to flying around in close proximity to the aircraft when released, but with careful modulation of the ERUs, clean separation was demonstrated. Table 6 gives some idea of stores configurations that have been cleared for the Hawk, up to about 1981.

Table 6 External Stores (1981)

STORE TYPE PYLON STATION

Under Inboard Outboard Fuselage Wing Wing

Aden 30 mm Gun Pod X X X

BR 125 125 kg bomb (free fall) X X X

BR 250 250 kg bomb (free fall) X X & XX X & XX

BRP 250 250 kg bomb (retarded) -- X X

BR 500 500 kg bomb (free fall)* X X X

Mk.81 250 lb bomb (free fall) X X X

Mk.81 SE 250 lb bomb (retarded) -- X X

Mk.82 500 lb bomb (free fall) X X & XX X & XX

Mk.82 SE 500 lb bomb (retarded) -- X X

Mk.83 1000 lb bomb (free fall)* X X X

540 lb MC bomb (free fall) X X X

BL 755 Cluster bomb X X X

Matra F155 M/N Rocket launcher -- X X

Matra F2 Rocket launcher -- X X

LAU 51 Rocket launcher -- X X

Oerlikon Snora Rocket launcher -- X X

CBLS 100 Practice bomb carrier** -- X X

CBLS 200 Practice bomb carrier** -- X X

455 litre (100 IG) External fuel tank -- X --



Sidewinder AIM-9G -- X --

Matra Magic Air-air missile -- X --

Reconnaissance camera pod X -- --

Sea Eagle Air-surface missile X -- --

All are 14 inch twin suspension stores. X indicates single carriage

* Modified trailing edges to fins. XX indicates carriage on twin store carriers

** To carry free fall and retarded practice bombs


46

The first Mk.50 series orders were from Finland, after a delegation from their Air Force paid a prolonged visit to Kingston, looking at every aspect of the aircraft in detail, and putting in a number of flying hours. They ordered 50 aircraft, of which all but the first 4 were to be assembled in Finland by the Valmet company. The contract was signed in December 1977 and the first deliveries were made 3 years later. Perhaps the most voluminous of the stores arrays was a configuration of 8 x 250 kg free fall and retarded bombs on twin store carriers, with the gun on the centreline position (Figure 26). Sidewinder and Matra Magic missiles were cleared for use, and maritime possibilities were opened up by the carriage of “Sea Eagle” (Figure 27), and with the addition of 190 imperial gallon external fuel tanks. Neither of these were cleared for any customer. So many different arrangements of stores were flown that G-Hawk was informally dubbed ‘The Heinz jet’ – 57 varieties of stores! With these changes the Mk.50 series had a 30 % increase in take-off weight over the Mk.1, with 70 % more disposable load and 30 % more ferry range. Figure 27 Hawk ZA 101 with two AMRAAM, Subsequent orders (1980) two 130 imperial gallon tanks and Sea Eagle were from Kenya for 12 aircraft, and 20 for Indonesia.

Figure 26 Hawk Mk.60 series with heavy stores load Eight 250 kg bombs and gun pod


47

3.3 The Mk.60 series. This development, following the Mk.50 above, used the new wing dressing and also provided an additional ¾ flap position, to provide extra lift for take-off, but in general the 60 series did not have combat flaps. The Mk.67 did have a combat flap setting, as well as a long nose and nosewheel steering, but no laser or FLIR. Adaptive anti-skid wheel brakes were fitted, again to improve wet runway performance. The wheels and tyres were further upgraded. Take off, acceleration and sustained turn performance were all enhanced by fitting the Adour Mk.861 having an uprated 5,700 lb SLST. This was 10 % higher at take-off, but the maximum rating in flight gave some 20 % increase in thrust over the Mk.851 at sea level, at 0.8 Mach no, as shown in Figure 28.

Figure 28 Adour performance - maximum rated thrust, sea level, ISA conditions Underwing tanks of either 130 or 190 imperial gallon capacity were cleared, the latter for ferry cases. With the extra power, the maximum take-off weight could be raised by about 17 % over the Mk.50, and the total disposable load by a third (corresponding to 50 % and 125 % respectively higher than the T.Mk.1, with the ferry range up by 65 %). The opportunity was taken to fit a more sophisticated weapons management system and to embody structural improvements designed to increase fatigue life and further simplify build and maintenance An order for 8 aircraft was received from Zimbabwe, the first delivery being in July 1982. This order had a sad start. Within days of five aircraft arriving at Gweru airport, they were attacked at night by saboteurs with timed explosive charges. One aircraft was destroyed and the other


48

four badly damaged, though they were eventually repaired. Ten years later Zimbabwe placed a second order, for five Mk.60A aircraft. Subsequently, orders were received from Dubai (9 aircraft), Kuwait (12 aircraft), Abu Dhabi (16 aircraft), Saudi Arabia (30 aircraft) and Switzerland (20 aircraft for delivery in 1990). 3.4 The T-45A Goshawk (Covered in detail in PART 2) After a competition with several contenders, the U.S. Navy selected a version of the Hawk, modified for carrier operations, as part of their VTXTS training system, in 1981. Hawker Siddeley’s leading partner in this venture was Douglas Aircraft Company of Long Beach, and long a supplier of USN aircraft. They were part of the McDonnell Douglas group, with which HSA already had links via the Harrier (AV-8 in the USA). The potential order was for over 300 aircraft, the largest order received. Later, ironically, when the USN discovered how much training the Hawk could pack in to a flight, the numbers were reduced! Many modifications were needed, including strengthened, long stroke undercarriage, twin nosewheels compatible with catapult launch, tail hook, side-mounted airbrakes, and the full flap vane with Tailplane Canard Vanes, called “SMURFs” in the USN. Initially a version of the Adour Mk.861, de-rated to 5,450 lb SLST, was fitted, but later changed. Some early aircraft had a plain leading edge, but leading edge flaps were developed and these were retrofitted to earlier aircraft. The fin height was increased by 6 inches greater than the other versions of the Hawk. The big problem was to obtain a high value of lift with acceptable handling, but this was eventually settled (see Part 2). Figure 29 shows some of the external features, in a full flap carrier approach.

Figure 29 U.S. Navy T-45 Goshawk Note the external changes from a

standard Hawk: full span leading edge slats with breaker strip inboard Strengthened and longer stroke undercarriage, twin nosewheels, arrester hook

and side-mounted airbrakes, tailplane canard vanes (‘SMURF’s in U.S.N.)


49

3.5 The Mk.100 series. The Mk.100 version (Figure 30) differed quite markedly from the original, particularly in terms of equipment and a longer nose. The wing now had permanent wing tip installations for air-air missiles and a fixed droop was added to the leading edge of the wing to enhance lift further, about 20 %, in the Mach 0.7 region. Most of the previous improvements of the Mk.60 series were retained.

Figure 30 Series 100, production layout, with pods and air-air missiles The full flap vane was restored, and the Tailplane Canard Vanes were installed. Notice the small size of the tailplane canard vanes, just ahead of the tailplane. The biggest changes were in the avionic systems. This now included optical laser ranging and FLIR sensors for navigation and attack, and an inertial navigation system. An advanced Head Up Display and a Weapon Aiming Computer was introduced, with upgraded cockpit displays (‘glass cockpit’). All this operated via a MIL 1553B databus. The powerplant now was an uprated version of the Adour, the Mk.871, which had some 15% more thrust at temperate conditions, and more than 25 % more, “hot and high”. Other changes included HOTAS (Hands On Throttle And Stick). There was no provision for an ECM (Electronic Counter Measures) pod, although one was flown. The maximum external load was raised to 7,200 lb. A three-view G.A and a cut-away drawing are shown in Figures 31 and 32.


50

Figure 31 Hawk 100 three-view G.A.

Note the tailplane canard vanes and tip-mounted air-air missiles

Figure 32 Hawk 100 cutaway drawing

Nose has provision for FLIR and or laser sensors Some aircraft have provision for installing an air refuelling probe

The fin incorporated a Radar Warning Receiver in the leading edge.


51

The hard-working G-HAWK was converted to the new configuration and made its first flight in this form on 1st October 1987. In the early 90’s orders for the 100 series had been received from Indonesia (8 aircraft), Abu Dhabi (18 aircraft), Malaysia (10 aircraft) and Oman (4 aircraft). 3.6 The Mk.200 series (single-seater). The Mk.200 version has elements of both the Mk.60 and Mk.100 series and is quite a potent attack aircraft (Figure 33).

Figure 33 Foreground: ZH 101 as series 100 development aircraft Background: Hawk series 200 single seat combat aircraft

Both aircraft have fin-mounted Radar Warning Receiver and enlarged tail cone The main changes to the airframe were in the deletion of the front pilot position and the re-siting of the nosewheel, but something like 80% of the airframe is common with the Hawk 100 series. The nose mounted radar necessitated a change from the nose pitot-static system of the two-seaters to one having twin compensated sensors on either side of the fuselage. Heavy stores can be carried, the maximum store load being approximately 6,800 lb. The series 200 has an AC generation system, and a wide range of different avionic systems (similar to those of the series 100) was offered. Included in the armament fit were twin 30 mm Aden Mk.4 cannon, with the option of twin 27 mm Mauser cannon permanently mounted in the fuselage. This freed the centreline pylon position to carry a 130 imperial gallon fuel tank if required, though in the end, the centreline pylon was not equipped to pass fuel. Whilst undergoing development at Brough, the internal gun system considered was 2 x 20 mm cannons, then reduced to one. Finally, no internal gun was fitted. Of course, more power is needed with these heavy loads, and the uprated Adour Mk.871 is fitted, like the 100 series. The Martin Baker ejector seat is the Type Mk.10L.


52

A Fairey Hydraulics yaw control system is fitted to a rudder actuator and servo control is via an autostabilisation computer. Figures 34 and 35 show a three-view general arrangement of the aircraft and a cutaway drawing respectively.

Figure 34 Hawk 200 three-view G.A. Initial configuration, based on Mk.60/100 series

Aircraft common with two-seater aft of cockpit rear bulkhead

Figure 35 Hawk 200 cutaway drawing


53

The first prototype single seat aircraft (ZG 200) made its first flight on 19th May 1986, flown by Mike Snelling. Unfortunately the aircraft was lost on its 43rd flight on 2nd July 1986, in an accident which was attributed to g-induced loss of consciousness by the pilot, Jim Hawkins, who very sadly lost his life. The first pre-production Hawk 200 (ZH 200) flew on 24th April 1987 flown by Chris Roberts. A third demonstrator Series 200RDA (ZJ 201) flew on 13th February 1992, equipped with full avionics and systems, and with Westinghouse AN/APG-66H radar. The first production aircraft (for Oman) flew on 11th September 1993, and the first for Malaysia on 4th April 1994. As of 1995, orders had been received from Oman (12 aircraft), Malaysia (18 aircraft) and Indonesia (16 aircraft). See Table 8 in section 3.8 for an update on numbers. 3.7 Comparision of Hawk variants. This is shown in summary on Table 8. It shows how the Hawk evolved over the years, as the need arose.

Jour

nal o

f Aer

onau

tical

His

tory

Pap

er N

o. 2

013/

01

54

Tab

le 7

Haw

k D

evel

opm

ents

(up

to

1995

)

T

.Mk.

1

Mk.

50Se

ries

Mk.

60 S

erie

s

Mk.

100

Seri

es

Mk.

200

Seri

es

Ado

ur E

ngin

e T

ype

M

k.15

1

Mk.

851

M

k..8

61

M

k.87

1

Mk.

871

Nom

inal

SL

ST, I

SA

(lb)

5,2

00

5

,340

5,7

00

5

,845

5,8

45 (

6,03

0)

(N)

23

,130

23,7

50

25

,350

26,0

00

26

,000

(27

,000

) D

imen

sion

s

Win

g sp

an

(m

)

9.39

9.

39

9.39

9.

94

9.94

(f

t)

30

.8

30.8

30

.8

32.6

32

.6

W

ing

area

(m2 )

16

.69

16.6

9

16

.69

16.6

9

16

.69

(ft2 )

17

9.6

179.

6

17

9.6

179.

6

17

9.6

O

vera

ll le

ngth

(m

)

11.8

4

11

.84

12.4

3(M

k.67

)

12.4

3

11

.35

(ft)

38.9

38

.9

40.8

40

.8

37.2

5

T

ailp

lane

are

a

(m2 )

4.

328

4.32

8

4.

328

4.32

8 (e

x T

CV

) 4.

328

(ex

TC

V)

(ft2 )

46

.6

46.6

46

.6

46.6

46

.6

Fi

n ar

ea

(m

2 )

2.50

8

2.

508

2.50

8

2.

61

2.61

(f

t2 )

27

27

27

28.1

28

.1

Wei

ghts

E

mpt

y

(kg)

3,62

2

4,01

2

4,

400

4,45

0

(l

b)

7,

986

8,

845

9,70

0

9,

810

Max

imum

wea

pon

load

(k

g)

1,

500

3,00

0

3,

000

3,00

0

3,

000

(lb)

3,30

0

6,

613

6,61

3

6,

613

6,61

3

Jour

nal o

f Aer

onau

tical

His

tory

Pap

er N

o. 2

013/

01

55

Tab

le 7

(co

ntin

ued)

H

awk

Dev

elop

men

ts (

up t

o 19

95)

T.M

k.1

M

k.50

Seri

es

M

k.60

Ser

ies

M

k.10

0 Se

ries

Mk.

200

Seri

es

Wei

ghts

(co

ntin

ued)

Max

imum

fue

l (in

t.)

(k

g)

1,

346

1,34

6

1,

304

1,30

4

1,

304

(s.g

. 0.7

9)

(l

b)

2,

963

2,96

3

2,

876

2,87

6

2,

876

(

ext.)

(kg)

--

93

2

932

93

2 (1

,360

fer

ry)

93

2 (1

,360

fer

ry)

(lb)

--

2,05

5

2,

055

2,05

5 (3

,000

fer

ry)

2,05

5 (3

,000

fer

ry)

Max

imum

take

off

wei

ght

(kg)

5,70

0

7,

350

9,10

0

9,

100

9,10

0

(l

b)

1

2,57

0

1

6,20

0

2

0,06

0

2

0,06

0

2

0,06

0 P

erfo

rman

ce

Max

. lev

el s

peed

(kno

ts)

535

5

35

545

540

540

(k

m/h

our)

9

90

990

1,

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1,00

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1

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M

ax. l

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Mac

h no

.

0.8

7

0

.88

0.88

0.

88

0.

88

Serv

ice

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ling

(m

)

14,6

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15,2

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1

4,00

0

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3,70

0

(f

t)

48

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46,

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4

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ith

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56

Tab

le 7

(co

ntin

ued)

H

awk

Dev

elop

men

ts (

up t

o 19

95)

T.M

k.1

M

k.50

Seri

es

M

k.60

Ser

ies

M

k.10

0 Se

ries

Mk.

200

Seri

es

Nor

mal

sym

met

ric

‘g’

60%

fue

l, 1,

360

kg, 3

,000

lb s

tore

s

+8

-4 (

1,50

0 lb

) +

8 -4

+

8 -4

+

8 -4

+

8 -4

60%

fue

l, 2,

720

kg, 6

,000

lb s

tore

s

---

+

6 -3

+

6 -3

+

6 -3

+

6 -3

M

axim

um in

stan

tane

ous

turn

s 60

% in

t. fu

el +

2 o

r 4

Side

win

ders

Sea

Lev

el

(

2 S/

W)

+ 7

.5 g

(4

S/W

) +

7 g

2

0,00

0 ft

+ 4

g

+ 3

.8 g

M

axim

um s

usta

ined

tur

ns

60%

int.

fuel

, cle

an

Sea

Lev

el

+

5.8

g

+

5.9

g

20

,000

ft

+

3 g

+ 3

.8 g

60%

int.

fuel

+ 4

Sid

ewin

ders

, Sea

Lev

el

+3.

9 g

+

3.9

g

20,

000

ft

+ 2

g

+

2 g

C

omba

t ra

dius

(nm

) w

ith g

un p

od, 2

xAIM

9, 4

x 5

00lb

bom

bs, 2

x 1

30 im

peri

al g

allo

n ex

tern

al ta

nks

345

A

s ab

ove

but w

ith 4

x 1

000

lb b

ombs

1

25

Fer

ry r

ange

(nm

) C

lean

1,

313

W

ith 2

x 1

30 I

G e

xt. t

anks

1,57

5

1,

360

1,36

5

W

ith 2

x 1

90 I

G e

xt. t

anks

(fl

own

but n

ot c

lear

ed)

1,60

0+

!,

400+

1,40

0+


57

3.8 Total Hawk orders and deliveries. These are shown in Table 8, collated from various sources, but mainly from Janes “All the World’s Aircraft”. It is updated to reflect the position as in 2011.

Table 8 Hawk Orders and Deliveries (as of 2012) Country Type Number Dates Contract (C) or delivery Abu Dhabi Mk. 63 16 October 1984 to May 1985 (14 a/c converted to 63A and 2 of these to 63B) Mk. 63C 4 February 1995 to March 1995 Mk. 102 18 April 1993 Australia Mk. 127 34 April 2000 to October 2001

(including 1 test airframe) Bahrain Mk. 129 6 October 2006 to December 2006 Canada Mk. 115 23 July 2000 to August 2004

(1 to Oman) Dubai Mk. 61, 8 March 1983 to September 1983 Mk. 61A 1 June 1988 Finland Mk. 51 50 December 1980 to October 1985 Mk. 51A 7 November 1993 to September 1994

(Mk. 66 16 Purchased from Switzerland, not included in total)

Table 8 (continued) Hawk Orders and Deliveries (as of 2012)

India Mk. 132 24 November 2007 to November 2008 Mk. 132 42 August 2008 onward Mk. 132 40 2010 (C) Mk. 132 17 Indian Navy Indonesia Mk. 53 20 September 1980 to March 1984 Mk. 109 8 May 1996 to March 1997 Mk. 209 16 February1996 to March 1997 Mk. 209 16 April 1999 Kenya Mk. 52 12 April 1980 to May 1982 Kuwait Mk. 64 12 November 1985 to September 1986 Malaysia Mk. 108 10 January 1994 to September 1995 Mk. 208 18 August 1994 to May 1995 Oman Mk. 103 4 December 1993 to January 1994 Mk. 203 12 December 1994 to May 1995 Saudi Arabia Mk. 65 30 August 1987 to October 1988 Mk. 65A 20 March 1997 to December 1997 Mk.165 22 Contract 23 May 2012 (RAF T.Mk.2 standard) South Korea Mk. 67 20 September 1992 to August 1993 South Africa Mk. 120/LIFT 24 May 2006 to 2008 Switzerland Mk. 66 20 November 1989 to November 1991

(WDS 2002)


58

United Kingdom T.Mk. 1 176** November 1976 to February 1982 (** Contract document 175) Mk. 128 28 April 2009 onward (T.Mk.2) United States* T-45A/C 223 April 1988 to October 2009-

(incl.2 prototypes) (* with McDonnell Douglas, now Boeing) Zimbabwe Mk. 60 8 July 1982 to October 1982 Mk. 60A 5 June 1992 to September 1992 TOTAL 994 Demonstrators, etc. G-HAWK, ZA 101 Mk. 50 1 Rebuilt with Mk.102 extended nose. ZJ 100 Mk. 100D 1 ZJ 951 Mk.100 NDA 1 New Development Aircraft ZG 200, ZH 200, Mk. 200 2 ZJ 201 RDA Mk. 200 1 Radar Development Aircraft TOTAL 1000

Data from Jane’s All the Worlds Aircraft 2011-12 Edition, plus communications with R. Storey, C. Hodson and C Farara.

3.9 Future Hawk development and sales. It is quite remarkable how the original Hawk airframe has been so successfully developed and equipped with continually updated equipment. There is at least one Hawk flying in a research role with software giving variable stability and there may be further developments in this area. Many developments have occurred after 1995, which is where this part of the story ends. Orders have been received from Australia (Lead-in fighter) and Canada with enhanced capability. This part of the Hawk story is covered in Part 3. What is certain is that efforts will be made to keep the Hawk viable in today’s environment, and that they will succeed.


59

4. The Rolls-Royce Turbomeca (RRTM) Adour turbofan engine. 4.1 A brief description of the Adour engine. The Adour is a 2 shaft bypass jet engine having a bypass ratio of approximately 0.8, though this varied as the engine was developed. The first stage of the two stage fan (with a transonic tip speed) is made of titanium, the second being aluminium alloy. The fan is driven by a low pressure turbine. The five stage HP compressor rotors are also made from titanium, with steel stators, the compressor being driven by a single stage HP turbine having air cooled blades. The overall compression ratio is 11. The combustion chamber is single annular unit, with 18 fuel injector nozzles. The “dressed” weight of the engine is 1,307 lb (593 kg) and its overall length is 76.9 inches (1,947 mm). The intake diameter is 22.3 inches (867 mm). Figure 36 is a cut-away drawing of the engine, and shows the three-point mounting. It is of modular construction, with large assemblies which can be removed and replaced in one piece, saving servicing time at the engine.

Figure 36 Rolls-Royce Turbomeca Adour engine The engine was first used in a re-heat configuration in the Anglo-French SEPECAT Jaguar twin-engined aircraft, and also in the Mitsubishi F.1. For the Hawk T.Mk.1 engine, the Mk.151, the re-heat tail pipe and equipment were replaced by a normal jet pipe, and a simpler control system. This adaptation from a known engine saved weight and cost, and cut risk. There was 95 % commonality between the two engines. The engine was Government furnished equipment, so that negotiations on cost, etc. were between Rolls-Royce and MoD. All the uprated variants of the basic engine were similar in build, with changes of material, etc., and all may be operated with AVTAG or AVTUR fuel.


60

4.2 Performance of the engines. The export engine, the Mk. 851, was essentially the same as the engine of the T.Mk.1, the Mk.151. The nominal thrust at take–off rating for ISA Sea Level Static (SLST) was quoted as 5,200 lb for the 151, but this was a minimum guaranteed value. For the export aircraft it was quoted as 5,340 lb. There were life improvements and improved cooling of the turbine blades and stators. An acceleration switch and enhanced fuel pre-heating were incorporated into the control system. The first uprating (Adour Mk. 861), was for the Mk.60 series Hawks, giving about 5,700 lb SLST, ISA, but this increment increased with forward speed (see Figure 28). This was achieved by increased turbine entry temperature and re-matching. It was also T2 compensated. The fan aerodynamics were improved and the HP turbine blades had revised aerofoils and improved cooling. The material of the LP blades was improved. There were changes to the control system, including an increased capacity fuel pump. The second uprating (Adour Mk. 871) was for the series 100 Hawks and the single seat variant, the series 200. Here the SLST was quoted as 5,990 lb in a 1992 brochure, with an increase of nearly 50% over the Mk.151/851 at M=0.9 at sea level. Modifications included titanium stage 1 stator blades and stage 2 rotor blades, INCO 718 discs for both HP and LP turbines and improved blade tip sealing, cast directionally solidified blades for the HP turbine, and cast single crystal blades for the LP turbine. There were improvements to the combustor, and changes to reduce smoke. An LP speed limiter (NL / T) was introduced into the control system. This engine was also the F 405 for the USN, and it included a modified acceleration control which achieved 95 % maximum thrust only 3 seconds after throttle movement from the approach condition (70 % HP rpm). This compared with the standard engine which took 5 seconds to reach 95 % max thrust, though this was from ground idle at 55 % rpm. A ‘growth’ version, the Mk.881, was offered but as far as the author is aware, has not yet been used in a Hawk. This had an SLST of 6,300 lbf at max. rating, static. The latest version of the engine is the Mk.951 – details are given in Part 3. An idea of the performance of these differing Marks of engine is given by the data below, and Figure 28.


61

ADOUR Mk. 151/851 861 871 881** Thrust at Max Rating, static, uninstalled average engine, no bleed or off-take.

SL, ISA 5,240 5,710 5,990 6,300 (lbf)

SL, ISA + 24˚C 4,510 4,840 5,500 ------ (lbf)

SL, ISA + 35˚C 4,160 4,400 5,120 5,850 (lbf) At SL, ISA conditions.

Air mass flow 94.0 94.7 97.6 101.9 (lb/sec)

Bypass ratio 0.79 0.78 0.76 0.73

Overall pressure ratio 10.7 11.2 11.3 11.5

N1/N2 99.7/100.7 102.5/101.5 108/100.2 101.7/101.9 (%)

(100 % N1 = 13,600 rpm, 100 % N2 = 15,512 rpm.)

SFC 0.71 0.74 0.78 0.76 (lb/hr/lb.th) At M = 0.8, ISA, Sea Level, Maximum Rating.

Thrust 4,150 4,760 5,760 6,680 (lbf)

SFC 1.06 1.07 1.08 ------ (lb/hr/lb.th) ** It is unclear if this variant was actually built, but ultimately it led to the 951. This newest variant of the Adour, the Mk.951, is discussed in Part 3.


62

5 CONCLUDING REMARKS FOR PART 1 Why was the Hawk programme so successful? The author, in his personal opinion, believes this was due to a number of factors: There was extensive research of the market and requirements.

There was a thorough understanding of the operational tasks, and a conviction that developed strike versions were essential.

Cost estimates were devised early on so that reasonably firm estimates could be made of the effect of design changes, and the results were passed on to the design and production engineers.

Risk was reduced by the maximum use of known technology, and by early wind tunnel tests to highlight areas of improvement at an early stage.

An existing engine was used, so that no long periods of testing were needed to prove the engine.

It was not an international collaborative programme

All aircraft were built on production tooling. There were no prototypes as such, and there was one contract to start with – no batch orders – for a substantial home order which was a sound footing on which to expand.

There was rigid project control, and excellent liaison between MoD, the operator and the company.

There were talented, enthusiastic and dedicated engineers at all levels, all believing in the success of ‘their’ aircraft. At the Hawker Siddeley works and airfield at Dunsfold, staff in Flight Test and the assembly of production aircraft made huge efforts. Figure 37 shows the complex there, in tribute to them. Figure 38 shows some of the ‘Hawk People’ in a group photograph taken at Dunsfold in 1995, on the 21st anniversary of the first flight of a Hawk, XX 154, which is seen in the background, and which was reported to be still in use at the A. & A.E.E., Boscombe Down, early in 2012, nearly thirty-eight years after it made its first flight.


63

Figure 42 Hawk ZA 101 carrying multiple rocket launchers overflies HSA (Kingston) Airfield and Works at Dunsfold

Figure 43 Hawk 21st Anniversary of first flight Some Hawk people SOURCES OF INFORMATION

HSA Kingston publication HSK 27 “The HS 1182” October 1970 Air Staff Requirement ASR 397 (2nd Draft) HSA publication HSK 30 “HS 1182 Choice of Power plant.” January 1971


64

HSA publication HSK 51D Options of the HS 1182 July 1971 Ministry of defence (Procurement Executive) Specification 281 Development and Production March 1972 RRTM 18 Rolls-Royce Turbomeca RT 172-06 Brochure June 1973 RRTM 151 Rolls-Royce Turbomeca Adour Mk.871 / F 405 April 1992 Flight Logs – Various Aircraft 1974 – 1989 A&AEE Letter Report Hawk T.Mk.1 March 1976 Performance & Handling Report for C.A. Release Also April 1976 KGT.R.00386 “Hawk 1/3 scale intake static tests.” July 1976 HSA publication HSK 193 “H.S.Hawk GA/TR Aircraft Brochure” March 1977 BAe(K) 90 “Single seat Hawk – An Appraisal of Possible Variants” November 1979 BAe(K) 376 “Technical Description, Hawk 200 series” June 1983 BAe(K) 534 “Technical Description, Hawk 60 series” 1985 AKN.SEG.006 “Hawk 100 series Technical Features” April 1986 Hawk 21st Anniversary Booklet November 1995 Hawker Association Newsletter, No.31. Autumn 2011. Janes “All the Worlds Aircraft” Several Editions. PowerPoint presentations at R.Ae.S Headquarters. 20th October 2011 NATOPS Flight Manual Navy Model T-45A 15th January 1997 “British Aerospace Hawk into the 1990s” G. Chisnall Proceedings of the Institution of Mechanical Engineers, London Vol. 206 1992 “Hawk Comes of Age.” Peter R March RAF Benevolent Fund ISBN1-899808-00-0 1995 Private Communications 2011, 2012 from R. Storey, Engineering Manager, Aerodynamics, BAE Systems Ltd., Brough. Brochure “Hawk, Advanced Jet Trainer”. BAE SYSTEMS PLC 2010.

APPENDICES

APPENDIX 1.1 (a) HAWK T.Mk.1 WIND TUNNEL MODELS The letters heading this list of wind tunnel models refers to the descriptions of the models in section 2.3.2, page 23.


65

Low Speed Time Scale (a) 1/2 scale half model for V/STOL 15ft tunnel at Hatfield (Preliminary configuration) 1970 / 71 (a) 9/16th scale updated version of (a) 1971 / 75 (d) 1/10th scale preliminary complete model for Woodford 1971 / 72 (h) 1/3rd scale intake + nose fuselage + stub wing for duct testing Updated later to low intake configuration, Kingston and Woodford 1971 / 74 (f) 1/6th scale definitive complete model for Woodford 1973 / 75 1/6th scale stores for testing with (f) 1973 (i) 1/18th scale preliminary spin tunnel model (IMF, Lille) 1972 (i) 1/18th scale definitive spin tunnel model, moving controls (IMF, Lille) 1973 (j) 0.315 scale full model for the RAE (Farnborough) 5m pressurised low speed tunnel, also used in the 13x9ft low speed wind tunnel at Weybridge ca.1979 / 80 High Speed (b) 1/30th scale preliminary model and modifications, tested in High Speed “Blow-Down” tunnel at Brough 1971 / 72 (b) Model (a) tested in the RAE 3 ft high speed wind tunnel 1972 (b) 1/28.5th scale updated model (a) tested at Brough and rebuilt later for store release testing 1973 + (e) 2-diml. model aerofoil tests, 3 models and mods. ARA Bedford 1971 / 72 (g) 1/6th scale definitive model, including plain and pressure plot wings, And pressure plot canopy, ARA Bedford test to M=0.9 1973 / 74 1/6th scale intake and duct performance – new fuselage for (g) 1973 (g) 1/6th scale rear fuselage modification for airbrake tests and extension to M = 1.2 . ARA Bedford 1975 / 76 (g) 1/6th scale RAF stores for (e) including store and pylon loads 1974 Total cost of model manufacture was approximately £ 160,000 at then prices, with the exception of (j) which was jointly funded by BAE and MoD for research into stalling characteristics, particularly with pylons and stores mounted on the wing. Total cost of testing (with the exception of (j)) was approximately £250,000 at then prices. Other, later, models were a 1/13.9 scale half model for Brough, and a new 1/14.6 scale high speed model for the Warton 1.2 m HST.


66

APPENDIX 1.1 (b) HAWK T.Mk.1 FLIGHT TESTING HOURS. Four development aircraft at BAe. Two at A&AEE. XX 154 Handling, Flutter, Loads XX 159 XX 156 Systems and Engine XX 160 XX 157 Avionics and Weapons XX 158 Stall, Spin, Performance Total flight test hours to C.A. Release 600 to 650 500 to 550 Post C.A. Release, mods. etc. 150 approx. 130 approx TOTAL 800 approx 700 approx SIGNIFICANT DATES 1st issue of HSA internal draft Spec. November 1968 Official go-ahead October 1971 Specification agreed March 1972 First Flights XX154 August 1974 XX156 May 1975 XX 157 April 1975 XX 158 July 1975 XX 159 June 1975 XX 160 November 1975 Initial C.A. Release (Flying Trainer) October 1976 2nd C.A. Release (Weapon Trainer) July 1977 1st delivery to RAF Valley November 1976 1st delivery to RAF Brawdy July 1977 1st Trainer RAF pilot solo July 1977 1st delivery, export aircraft (Finland) April 1980 By mid 1980, 140 aircraft had been delivered, and approximately 70,000 flying hours had been accumulated. The last RAF Hawk was delivered in 1982. The RAF fleet, then of 132 Hawks, achieved 1 million flying hours on 5th July 2006.


67

APPENDIX 1.2 COSTS HAWK CASE STUDY (a) ANNEX F to House of Commons Defence Committee Paper 145/2. HoC Defence Committee Paper 145/2 contains some interesting material on the order for the Hawk, particularly on the acquisition cost at entry into service, in January1971 pounds prices. Aircraft development was given as costing £22.2 million, with unit production costs of £437,000. It is not clear that this excluded the engine, which was a Government furnished item. A typical engine price might be £140,000, so if this were included, the unit fully equipped airframe cost would be about £300,000, which seems a bit low. If the above figures are added up, then the total cost for the 175 aircraft comes out to be £98.675 million. On top of this, the reliability and maintenance incentives totalled £2.5 million, and Hawkers won the lion’s share of this, so in round numbers, the whole Hawk fleet cost the exchequer only about £100 M, very good value even at January1971 prices. If the engines were separately purchased, say 200 (allowing for spares) at £140,000 each, this would add about £28 million to the bill, still good value. The document went on to give the timetable of events and some comments. December 1970 ASR 397 endorsed, calls for aircraft to enter service in 1976/77. (The first two Hawks were delivered in November 1976.)

May 1971 Initial Design Study contracts placed.

December 1971 Approval for launch of development and production.

February 1972 Treasury Approval for the above.

March 1972 Ministerial Approval. The aircraft evaluated were: BAC P.59, HS 1182 V, HS 1182 AT, 1182 AJ, Alphajet, Macchi 326 G, SAAB 105X and an improved Jet Provost. It was noted that the BAC P.59 was eliminated on cost grounds in September 1971, and that the Alphajet proposals for international collaboration were unattractive. It may be that the last two were eliminated because they had side-by-side seating. The Hawk met or exceeded all the ASR requirements with the exception of the threshold speeds, which were marginally higher. The Paper notes that there were special conditions for the Hawk contract which affected the cost, namely:

The Contractor had completed the basic design as a private Venture.

There was a low level of technology risk.

The engine was low risk.


68

It was a fixed price for 175 aircraft.

There were no Specification changes during development. There was also a comment that these favourable conditions were unlikely to arise in future contracts. (b) Hawk ‘War Role’ aircraft conversion costs. In another document dated 2nd July 1982, it was stated that the cost per aircraft of the conversion was £61,100 for 88 aircraft, a total of £5.38 million. This included a P.V. levy of £242,000, or £2,750 per aircraft. Mod kits ordered totalled £ 4.73 million. (c) Estimated cost of a Hawk combat aircraft for export. In May 1980, an estimate was compiled in house for a fully-equipped combat version of the Hawk, to be exported. The cost breakdown was as follows, per unit: £ Labour, 42,000 man hour at £14.5 per man hour 609,000

Raw material 90,000

Bought out parts 580,000

Engine 460,000

Warranties, etc. 50,000

Tech. Pubs. Etc. 62,500

P.V. recovery 20,000

MoD levy at 5% 145,000

Bank/ Export Credit etc. 174,000

Allowances 7½% 217,500

Margin 20% net, 17½% gross 481,000

Total 2.889,000

APPENDIX 1.3 The ‘Phantom Dive’ phenomenon and its cure – in detail. The Phantom Dive phenomenon was introduced and discussed briefly in section 2.3.5. Because interest has been expressed in the way in which the problem was tackled and solved, the whole story is related here in some detail for those wishing to pursue it. Soon after the condition was first encountered it was recognised that the tailplane was not producing enough downwards force (negative lift) to counter the nose down pitching moment from the large flaps in their fully down (50 o) position. When the undercarriage was deployed,


69

the main wheel legs spoiled the flow through the flap slots to some extent, and this reduced the flap pitching moment sufficiently for the tailplane to cope. Similarly, reducing the flap angle by a few degrees also reduced the flap pitching moment, but there was a loss of lift and drag, essential for landing. The solution adopted for the Hawk T.Mk.1 was to remove the outer portion of the flap vane of the double slotted flap, as shown in Figure A3.1. This was a quick and effective solution, but when combined with the use of a fence and ‘breaker strips’ for stall warning, lost about 5 knots of stall speed, though the behaviour was benign.

Figure A1.3.1 Hawk T.Mk.1 showing extent of cut-back of the flap vane (port shown)

While acceptable for the T.Mk.1, this loss of stall speed had to be recovered for the US Navy T-45 version and for the combat variants of the Hawk, since the largest possible CLmax was required, with acceptable handling. Thus the flap vane had to be restored to its original length, and a cure had to be found for the tailplane stall by some means. An extensive programme of tests was laid on in the 15ft x 15ft V/STOL wind tunnel at Hatfield, using the 9/16 scale half model of the Hawk. The Reynolds number for the tests was 2.4 x 106 based on mean wing chord, high enough to be confident that the results would be representative of full scale flight. It was confirmed that the tail was stalling on its lower surface by photographing the behaviour of tufts attached to the undersurface. The results are shown in Figure A3.2, which shows the progress of the stalled flow across the tail at various tailplane angles at a low angle of attack.

outboard

375 mm


70

Figure A1.3.2 Analysis of tuft behaviour on underside of tailplane

Flap fully down (50o). undercarriage up. Full flap vane. Clearly, the more nose down the tailplane angle (- T), the more stalled flow is seen, suggesting

the lift (downwards) is becoming limited. With T still short of the full back stick position of -15o,

virtually the whole tailplane lower surface is stalled, even at the angle of attack for level flight, about 0 deg at low speed with full flap. The problem is worse with negative angle of attack. Figure A3.3 shows the variation of pitching moment with angle of attack on the model with tail off and tail on for T at 0 o and -12 o settings. The length of the vertical ordinate between the

tail off and tail on curves is proportional to the downwards lift developed by the tailplane. cannot be raised and the result is a dive of increasing steepness. At a positive angle of attack the effectiveness of the tailplane returns to some extent, in that there is some nose-up pitch available with the stick well back. With the u/c down, (full lines), back stick gives a nose-up pitch at all angle of attack due mostly to the change of tail-off pitching moment – the u/c legs spoiling the flow through the slots on the

T is tailplane angle

Angle of attack T = 0 o

T = - 12 o Looking at the broken line curves, for full flap vane and undercarriage up, it is clear that at about angle of attack of -10 o or so, pulling the tail from neutral (0 o) to nearly full back (-12 o) does not increase the tailplane downwards force at all, and there is virtually no positive pitching moment to bring the nose up. This is the consequence of the tail stall, and the ‘Phantom Dive’, where the nose


71

flap. With the short vane, (chain lines), the tail-off nose down pitching moment is further reduced, enhancing the effect. This was the cure for the RAF Hawk T.Mk.1. The obvious necessity was to enhance the lift from the tail at the fairly high angles of attack of the local flow at the tailplane. This could be done by cambering the tail airfoil section and this was tried on the model, but with only limited success. The age-old remedy, dating from the early 1920’s, was to add a leading edge fixed slot – this was the solution selected for the F4 Phantom – but the extra drag of the open slot at high speed meant it was not an option on the Hawk. This was when the Harrier ‘LERX’ (leading edge root extensions) solution occurred to us, promoted by B.V. Pegram, who had used them to extend the Harrier’s wing lift to higher angles of attack. In that case the extensions were attached to the wing, but this would not do for the Hawk tail because the tailplane power operating motor would probably not be able to cope with the increased pitching moment. The author modestly claims he suggested that the vanes could be fixed on to the rear fuselage ahead of the tailplane at such a position that their trailing edges were aligned with the tailplane leading edges at its most negative position. It was suggested that the vortex cast by the leading edge of the vane under these conditions should be effective enough to clear up the flow over the inboard end of the tail and increase the lift appreciably at a high local angle of attack. This harked back to the author’s experience in the early 60’s with the Slender Delta Research Aircraft, the Handley Page HP 115. Some experimentation was needed to fix the size and angle of the vane. Setting the vane at a negative angle enhanced the effect, but gave rise to some extra drag in the cruise. The flat position shown was effective enough on the model and might be expected not to cause more than a small amount of friction drag in normal flight. This was confirmed in flight. Figure A3.4 shows the vane selected and its position relative to the tailplane, and the approximate extent of the area affected by the vortex. A small part of the improvement is due to lift on the vane itself, but mostly the extra lift is due to the suction under the region covered by the vortex and the clean-up of the flow inboard.

Figure A1.3.3 Pitching moment curves, flaps fully down Tail off (lower curves) and tail on (upper

curves) with tailplane angles of 0 o and -12 o


72

Figure A1.3.4 Installation of Tailplane Canard Vane at 0 o Tailplane shown in fully nose down position. Note the small clearance between vane and tailplane

leading edge. The broken area shows the approximate position of the vortex wake from the vane leading edge

At HSA Kingston, the devices were known a ‘Tailplane Canard Vanes’ (TCV), but in the US Navy they were known as ‘SMURF’ (Side MoUnted Rear Fins). This caused some hilarity in the home camp at first, since there were well-known children’s cartoon characters of that name, but the USN was rather serious about it and insisted that as far as they were concerned this was how the device was to be known. The sketches of Figure A3.5 show the effectiveness of the TCV at cleaning up the flow over the inboard part of the tailplane underside. Clearly, more lift is being developed. The success of the TCV is shown in the pitching moment curves of Figure A3.6 for the model with TCV on. This has a small favourable effect as shown by the tail off curves, since it acts like a very small fixed tail – it provides a small amount of favourable pitching moment and a small increase in stability. With tail and TCV on, there is ample tailplane authority with aft stick (- T) at all angles of

attack in the usable range.

Tailplane vane at 0 o


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After checks in the ARA High Speed wind tunnel with the 1/6 scale model at high subsonic speeds which showed no effects due to the vanes, further checks in the spinning tunnel at Lille University again showed little effect. Further checks were made at a higher Reynolds Number, close to that for full scale, in the 5 metre pressurised, low speed wind tunnel at RAE (Farnborough), which confirmed the effect of the TCV. A full scale flight test was now sanctioned. This was to be conducted on XX156, one of the fully instrumented Hawks used by the MoD in their testing. They readily lent the aircraft to us, only stipulating that it should be returned to its T.Mk.1 configuration when the tests were finished. The aircraft was fitted temporarily with a wing dressed to T-45 standards and with bevelled flat plates representing the TCVs. It was first flown without the TCV to confirm that in this condition, the ‘Phantom Dive’ was indeed present, and then flown again with the TCV fitted. Figure A3.7 and A3.8 show the test set-up on the aircraft, from the side and from the top. Figure A3.9 shows the traces for flights without (left hand side) and with the vane (right hand side). In each case the most sensitive conditions for ‘Phantom Dive’ was selected, nominally an indicated air speed (IAS) of about 130 knots with flaps fully down and undercarriage up, and the aircraft loaded to a forward centre of gravity.

Figure A1.3.5 Effect of TCV on tuft behaviour on underside of tailplane. The region behind the vane

is cleaned up by the vortex springing from its leading edge

T = - 11 o

T = - 14 o

T = - 8 o

Figure A1.3.6 Pitching moment curves, tail off and tail on, with TCV at 0 o (broken line

with TCV off, for comparison)

T = - 8 o

T = - 9 o

T = - 12 o

T = - 0 o


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Figure A1.3.7 Installation of TCV on Hawk XX 156 with full flap vane. Side view, tailplane in neutral position.

Figure A1.3.8 Installation of TCV on Hawk XX 156 with full flap vane. Top view, tailplane at full nose down position.

Note the small clearance between the TCV T.E. and the tailplane L.E.


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Figure A1.3.9 Aircraft flight instrumentation traces. XX 156 with full span flap vanes. Attempted tailplane stalls at 130 knots IAS, with and without TCV fitted.


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The technique adopted was to push the stick forward to achieve a nose down pitch angle of about 10 degrees, then pull it sharply back to the back stop ( T of -15 o) and maintain it to

bring the nose up, if possible. The behaviour without the vanes is shown on the left hand side, and the aircraft is seen to approach the manoeuvre at 130 knots IAS, tail at -5 o and in level flight, pitch attitude zero. The angle of attack is zero, with flap fully down, nominally at 50 o. At 3½ seconds the stick is pushed forward ( T = +5 o) and the pitch attitude starts to go nose down with the angle of attack reaching approximately -13 o. The IAS starts to increase in the dive. About one second later the stick is pulled hard back and held there for about 9 seconds (t=14 seconds), during which time the pitch attitude goes increasingly nose down and the IAS continues to increase. The angle of attack remains well negative; there is no sign of the nose coming up. This is the ‘Phantom Dive’. At 14 seconds, noticing that the IAS is approaching the full flap limiting speed, the pilot relaxed the stick off the back stop over a couple of seconds (t from 14 to 16 seconds) and the angle of attack does start coming back towards zero, but at 16 seconds the flap has to be retracted as the IAS has built up. The aircraft starts to level out, but some 2,000 ft. altitude has been lost in the 16 seconds it has taken for the manoeuvre to be performed. Looking now at the right hand side traces (vanes on), a similar approach is made, though at a slightly higher altitude, about 11,500 ft compared with 10,000 ft before. The nose down push occurs at t = 6 seconds and is followed by stick hard back at 7 seconds. The aircraft responds immediately, the angle of attack reaching +9 o even though the stick has been returned to its earlier condition for T = -5 o. Pitch attitude has returned to level flight,

more or less, and IAS has hardly changed. A “non-event”, and the “Phantom Dive” has been cured. The effect of fitting the vanes was so markedly positive that the test pilot who flew both flights commented that usually when a new ‘fix’ is flown, and one is lucky, one gets a small but positive improvement. In this case he had never before seen such a dramatic improvement in behaviour from such a relatively small and simple device. The addition of TCVs was a cheap and effective solution of the problem of tailplane under surface stall, and allowed further development of the wing to try to achieve the low stalling speeds demanded for carrier operation in the US Navy, and for combat versions of the aircraft carrying heavy store loads. This was not quite the end of the story, because in the VTX project, the question was again raised, and further tests were made on a configuration closer to the T-45, with its side-mounted airbrakes. This is covered in Part 2 of this document.


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APPENDIX 1.4 Initial High Speed Development of the Hawk (This is a slightly edited version of the paper for a lecture to the delegation of the Finnish Air Force visiting Hawkers after their order, they being the first export customer. It is dated July 1978.) During the design development of the project, work was put in hand on the design and construction of a model to be tested in the high speed wind tunnel at Brough. Since the tunnel size was 27 inches x 27 inches (0.69m x 0.69m), the model was necessarily a small one, and a scale of 1/30th was chosen. Its layout was generally similar to the HS1182 version (see Figures 11 & 12 in part 1). The model came out to just over 1 foot in span. The wing sections were derived from previous 2-dimensional sections developed and tested for the Harrier and followed the peaky-leading-edge philosophy. The sweep and thickness-chord ratios were chosen having regard to the high speed cruise Mach number of 0.8, the requirement at that time. Since the model was sting-mounted, some enlargement round the tailpipe was necessary. Testing the model started in October 1971 and continued through 1972, despite the fact that the configuration had changed somewhat, to the ‘AJ’ design, to gain some general indication of the characteristics. The Reynolds number of the tests was fairly low, about 1.36 x 106 based on the mean chord, at M=0.8, though the tests covered from M=0.45 to M=0.9. Flight Reynolds number was about 11.5 x 106 at representative top speed conditions at altitude. The initial results were not encouraging. Severe flow breakdown was indicated on the wing, at only just beyond M = 0.8, and since the specification was now demanding a top speed of M = 0.85, it was clear some re-design was required at the higher speeds. However, even at M = 0.45, the shape of the pitching moment against incidence curves left much to be desired. The tail-off curve was fairly linear, and showed a distinct nose-down (favourable) break at about 11 o incidence. But when the tailplane was added, a distinct ‘ride-up’ occurred at a lower incidence, about 10 o, and the whole curve was much less linear. At higher Mach numbers, say 0.8, the curves for the wing itself are much more non-linear, showing evidence of flow separations, and the tail-on curves were completely unacceptable, with a pronounced pitch-up at quite a low incidence. For some time the aerodynamicists had felt uneasy about the high tail position, so several different positions were tried out – a low one at the bottom of the fuselage, and an intermediate position between with both a flat and an anhedral tailplane. None of these were much better at low speed (Figure A1.4.1), but all were better than the original tailplane position at M = 0.8 (Figure A1.4.2). Even so, none of the curves were particularly good. To try to explain these characteristics, extensive flow visualisation trials were undertaken and the dynamic pressures were measured in the flow field in the region of the tailplane.


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1/30th scale preliminary model tests (HS 1182A)

Figure A1.4.1 Effect of different tailplanes on CM versus Incidence High intake, original tailplane/body fairing Mach number = 0.45


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Figure A1.4.2 Effect of different tailplanes on CM versus Incidence High intake, original tailplane/body fairing Mach number = 0.8

At low Mach number, there was evidence of a strong vortex flow on the upper surface of the intake cowl, due, it seemed, to a strong outflow from the channel between the cowl and the wing upper surface. The vortex streamed back just under the tailplane, and this was confirmed by the dynamic head pressure contours, as well as by tufting. At higher speeds, though the vortex was still present, a strong shock wave developed in the channel, with boundary layer separation and a loss of sweep on the shock front. This was suspected as being the cause of the flow breakdown in the wing flow at slightly above M = 0.8, and because of the shed wake interfering with the tail, the cause of the pitching moment problem.


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The dynamic pressure contours showed that the tailplane enters a low pressure region at about 10 o incidence, but the wing itself stalls at 11 o, giving an overall pitch down at 12 o. As the tailplane height is reduced, so the loss of stability due to the loss of tailplane effectiveness occurs at a lower incidence, the low and anhedral tailplanes being the worse in this respect (Figure A1.4.1). At the higher Mach number, there is a large area of low dynamic pressure at the tail due to the wake shed from the wing/body/cowl region. Although the low tail and the anhedral tail begin to enter this region at about 2 degrees of incidence – shown on the low tail curve as a sudden loss of stability starting at that incidence – the wake is not very pronounced at these lower incidences and the effects are fairly small. However, the high tailplane, and to a lesser extent the intermediate flat and anhedral tailplanes, enter the low pressure region at a much higher incidence (7 to 8 o) where the flow is much more severely retarded, so that larger changes in stability occur. To counteract these problems it was decided to attempt to improve the wing/body junction – not necessarily in a practical way – by filling in the channel which had been shown to give very poor flow at high Mach number. The modification to the wind tunnel model is shown in Figures A1.4.3A and A1.4.3B. In addition, the rear fuselage was made less tapered at the tailplane junction. Whilst these improved the pitching moment curves at low speed, the high tail still gave a pronounced pitch-up beyond about 9 o at M = 0.8. The intermediate flat tail was better and the anhedral best of all, and reasonably acceptable. (Figure A1.4.4). Of course the shape changes made on the model were impractible on the aircraft, but showed the way forward.


Figure A1.4.3A Modified wing/intake

fairing (shown shaded) Figure A1.4.3B Cross section of fuselage

showing intake/wing fairing


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Figure A1.4.4 Effect of Fuselage Modifications on CL v/s Incidence Mach number. = 0.8

Three major decisions were taken as a result of this testing: 1. To re-design the basic 2-dimensional aerofoil section to improve its high Mach number behaviour and to extend the wing capability to cater for the increased cruise Mach number now specified. It was proposed to do this by applying new knowledge about the development of supercritical flow round the nose of aerofoils, and also imparting a mild degree of ‘aft loading’. At the same time, it was necessary to increase sweep slightly for balance and lay-out reasons, which should also be helpful. 2. To position the tailplane as low as possible, the rear fuselage was bent down from its original almost horizontal upper profile, and the tailplane given anhedral; its root was located above the tailpipe. A flat tailplane positioned very low was rejected because it was less favourable structurally, and the lack of fuselage side area beneath it would be small. This would reduce the damping in a developed spin, a contracted requirement. 3. The intake was lowered to a position just ahead of and above the wing root. This meant it was possible to form the sides of the intakes to produce a shape which gave an excellent wing/intake side junction, maintaining the sweep of the wing isobars right up to the fuselage


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sides. A possible problem was that the lowered intake might be more prone to ingestion, particularly of spray from the nosewheel. This was checked from an analysis of spray patterns from other aircraft and found to be acceptable. Later tests, running the aircraft through standing water, confirmed this. These three major changes, coupled with the reduced size of the aircraft in its ‘AJ’ form, necessitated the construction of further models. The new aerofoil design was checked out satisfactorily at high subsonic speeds at the Aircraft Research Association (A.R.A.) Two-Dimensional Tunnel, and the Brough high speed model was rebuilt to incorporate the changes, the scale now coming out as 1/28.5. At the same time, work was put in hand to construct a 1/6th scale high speed model to be tested in the ARA Transonic Wind Tunnel, since doubts were raised as to the validity of boundary layer transition fixing effects on the new aft-loaded wing sections on such a small model at relatively low Reynolds number, when it was tested early in 1973. It was not possible to resolve these issues at the time on that model without a large amount of work, so it was decided to rely on the larger ARA model for high speed characteristics. The 1/28.5 scale model went on to be modified to use the accelerated model rig in the Brough High Speed wind tunnel for store release testing. 1/6th scale High Speed Model Testing at A.R.A. One of the first aims in testing was to establish the number and position of boundary layer transition bands on the wings which would be required to produce valid data over a range of Mach number and incidence. A characteristic of wing designs which have substantial aft loading is that at model test Reynolds number there can be an interaction, particularly at high Mach number, between a shock-induced separation and a premature rear separation. With forward transition, the boundary layer thickness is greater than at full scale Reynolds number, and this interaction is accentuated, leading to a shock position that is not as far aft as it would be at flight Reynolds number. Even at attached flow conditions the shock may be slightly too far forward. The error in shock position, and hence in values of CL and CM will vary both with

incidence and Mach number giving misleading measured stability characteristics. It is normal practice, therefore, to choose transition band positions further aft on the surface in order to obtain a more representative boundary layer thickness at the shock and at the trailing edge. Suitable positions must be chosen so that various criteria are satisfied in the CL – Mach number

range of interest, namely:-

(a) It must be far enough ahead of the shock at separation onset to give a turbulent boundary layer – shock interaction;

but (b) not so far aft that the calculated boundary layer thickness is less than at full scale Reynolds number, and toward transition; and

(c) it must not be so near the shock that, as separation occurs and the shock moves forward under the influence of this separation, the process is arrested too soon by a spurious


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interaction between the transition roughness band and shock position, with the shock hesitating just down stream of the band. The shock position for separation onset at a given Mach number moves aft with increasing Mach number and so the ideal roughness band position also has to move aft with Mach number. The limits of incidence and Mach number for which a particular roughness band position is suitable have to be determined by calculations of boundary layer thickness for a given transition position to check on (b) and oil flow tests to indicate the shock positions for (a) and (c). Oil flow tests were made with transition bands at 5 % and 40 % chord positions and Figure A1.4.5 shows the progress of the shock front at about 50 % semispan with changing incidence. It can be seen from the Figure that when the shock reaches a position about 12 % behind the transition band, the shock movement stops, indicating an interference with the steady flow progression. Hence any results obtained at higher incidences should be viewed with caution. Figure A1.4.5 also shows that a forward transition band causes a shock to form and start to move rather earlier than is natural, due to the relatively thicker boundary layer produced at these low model Reynolds numbers, boundary layer calculations for the full scale aircraft at predicted cruise conditions (M = 0.84, CL = 0.3) showed that a 25 % chord position should give the best correlation with flight at the higher Mach numbers, so it was decided to use a 5 % transition for tests up to and including M = 0.8 and 25 % for higher Mach numbers, with an overlap at M = 0.8. ‘Limiting Incidences’ were defined above which the results might be suspect in magnitude, though possibly still useful in indicating the character of the curves above that point. The problem is clearly shown in Figure A1.4.6, which shows very different longitudinal characteristics depending on the type of transition fixing. Taking the 25% as representative of flight, it was felt that these need to be improved in the region M = 0.85 and 0.90. (Note that because the specification called for a maximum Mach number of 0.9, funding was not provided for tests at higher speeds.) Therefore, a long series of tests were put in hand to investigate suitable arrays of vortex generators in the wind tunnel. It is interesting to note that the aircraft flew in a wing configuration very similar to the initial one tested in the ARA tunnel, without vortex generators or, of course, transition strips. Pitch characteristics were assessed as marginal at the higher Mach numbers, indicating some agreement with the tunnel results, though to a lesser degree of difficulty. The final vortex generator array developed for the model consisted of a front row of eight vortex generators spaced across 12.6 inches, starting 14 inches from the centreline and toed in at 10 o, leading edges on the 25 % chord line, coupled with a rear row over the same region, but centred on 65 % chord and starting 0.45 inches further outboard. These were toed out by 20 o. Each vortex generator was 0.2 inches high and 0.45 inches long, the leading edges being cut back to 60 degrees sweep at the leading edges. (All dimensions are 1/6 model scale.) The front array was chosen to modify the shock wave/ boundary layer interaction over the outer middle part of the upper wing surface, and the rear row to suppress a premature trailing


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edge separation due to the low Reynolds number of the model. This would not be expected to occur at full scale.

1/6th Scale High Speed Model Test (A.R.A.)

Figure A1.4.5 Wing shock positions with different transition bands 50 % semispan Mach number = 0.80 and 0.85


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Figure A1.4.6 Effect of transition band position Tailplane angle = -3 o Mach No. = 0.8 and 0.85

The improvement in sectional characteristics due to the vortex generators is shown on Figure A1.4.7 where the local CL for a significant change in trailing edge static pressure (an indicator

of the onset of separation) is shown as a function of Mach number at four spanwise stations. The improvement is considerable, and as a result of these results, the vortex generator array for flight was chosen as the front row only (less one vortex generator inboard), the full scale vortex generator height being slightly lower at 0.15 inches model scale, 0.9 inches full size.


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Mach Number

Figure A1.4.7 Effect of vortex generators on wing sectional onset of boundary

layer separations due to compressibility

Full line – No vortex generators Broken line – Vortex generators at 25% & 65% chord Station Location “R” Just o/b of wing/fillet intersection “C” 55% semispan “D” 69% semispan “E” 83% semispan When flown, the improvement in characteristics was confirmed. Some time later, when it was shown that the aircraft could relatively easily exceed M = 0.9, funds were released to permit tunnel testing up to M = 1.2. The model was brought up to the then current aerodynamic configuration with two breaker strips and an outboard fence on each wing, and a more representative rear fuselage, with airbrake and ventral strakes. The rear row of vortex generators was retained on the model of course, for the reasons given earlier.


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These higher speed tests showed that the reduced longitudinal stability still apparent on the model between Mach numbers of 0.85 and 0.95 at low incidence disappears at 0.98, though the tailplane effectiveness decreases at M = 1.1 (Figure A1.4.8). As an overall view, it seems fair to say that the aircraft’s flying qualities above M = 0.9 - the originally specified design dive speed – are degraded but acceptable. This is reflected in the clearance in the Flying Trainer Role to supersonic speeds. As far as the 1/6 scale model is concerned, the results obtained seem to parallel the behaviour found in flight, but the deficiencies found on the model are much more pronounced than on the aircraft in flight.


Figure A1.4.8 Pitching characteristics, current model. Fences, breaker strips and vortex generators as flight (with rear vortex generators on model)

Flow over the canopy. Av.P.970 requires that canopy loads should be measured. In the case of the Hawk it was agreed that the integrated pressures over the canopy measured on the 1/6 scale model could be used to estimate the airloads, with inertia and pressurization loads added as appropriate.


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Plots of the static pressure distribution were integrated up over a wide range of conditions. In no case was there any evidence of shocks forming over the canopy, and this seems to be confirmed by the fact that there is no significant change in noise level in flight between subsonic and supersonic speeds at similar values of indicated air speed.


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PART 2 THE HAWK FOR THE U.S. NAVY VTXTS AND THE T-45 GOSHAWK

SUMMARY The author was inspired to write this account by the necessarily brief but excellent presentation on the subject given by Chris Roberts on 20th October 2011. The background of the proposal to establish a new U.S. Navy pilot training scheme (VTXTS) is discussed and a history presented detailing how this progressed through conceptual studies, culminating in the choice of a naval version of the Hawk. This necessitated a large volume of new design work in adapting the aircraft, particularly for carrier operations, and some of the changes required are described. A critical criterion which became apparent quite early was the minimum speed for a powered approach. The aircraft, in its early wing configuration, struggled to meet the requirement and much time and effort were expended to this end. Eventually full-span leading edge slats were fitted to the aircraft. All this cost a considerable amount of money and the McDonnell Douglas Corporation (as the main Contractor) took the US Navy to court to recover the money allocated, but finally settled out of court, to the apparent satisfaction of MDC. A report is given of what the proceedings covered. An acquisition history is shown and comprehensive performance data presented, which also give a good insight into the performance of other marks of Hawk aircraft having similar engines. 6. BACKGROUND AND HISTORY 6.1 Background. In the 1970’s, the U.S. Navy began preliminary moves towards replacing their pilot training system, the aircraft elements of which were the Rockwell T-2C Buckeye, a straight-winged tandem two-seater aircraft with twin engines (Figure 44) and the single engine, delta winged Douglas TA-4J Skyhawk (Figure 45).

D.Barrow

Figure 44 Rockwell T-2C Buckeye Pensacola, Florida


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(twin Pratt & Whitney JT-15 turbojets)

D.Barrow

Figure 45 Douglas TA-4J Skyhawk VT-7 Squadron (single Pratt & Whitney JT-52 turbojet)

On 28th February 1975, NADC (Naval Air Development Center, Johnsville) was tasked by NASC (Naval Aviation Schools Command) with the ‘Initiation of Advanced Jet Trainer Formulation’. At the end of that year their conceptual studies had established the criteria which would be required to be met by the whole training system, comprising all ground school material, computer-based learning, simulators and aircraft. This was called ‘VTXTS’ (Naval Aviation Experimental Training System) and would be the basis for Industry studies. A particular point that became apparent was the critical nature of the carrier powered approach speed and its maximum acceptable value, VPA MIN, defined later, a value for which was initially suggested as 115 kt on a 90 o F day at sea level, though an even more difficult preferred value was given as 105 kt. In September 1977 a Request for Proposal (RFP) was issued to a number of U.S. manufacturers with a view to conducting in-depth design studies into the feasibility of the complete VTXTS system, and in particular, the aircraft element. Some six months later, contracts were awarded to Douglas Aircraft Co (DAC), Northrop, Vought and General Dynamics (Fort Worth) for conceptual studies. 6.2 Conceptual Studies Some idea of what came out of the studies on the aircraft side was gleaned some time later from discussions between engineers from DAC and BAe. They had submitted two designs, one single-engine (Rolls-Royce RB 401-7 of just over 5,500 lb Sea Level Static Thrust) and one twin-engine (two Pratt and Whitney JT15D-5 of 3,000 lb SLST each). The selected final designs had moderate sweep, about 24 o at quarter-chord, with supercritical wing sections with average thickness to chord ratios of about 14 %. The wings were tapered and their aspect ratios were 6 for the twin and 7.25 for the single. The wing area was about 180 sq.ft and it was equipped with slotted flaps having a ratio of flap chord to wing chord of 30%.


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A total production run of 350 aircraft was envisaged with a total acquisition cost of some $2 billion (1978) and an overall timescale of nine years from go-ahead. The production should be complete in about 5 years. Foreign aircraft companies had been excluded from this phase, but had been told that their companies would be considered in the next phase of studies. Members of DAC staff visited BAe in October and November 1978, but soon after confirmed that they would be working on their own. 6.3 Technology Base Study for existing aircraft. Contracts were awarded on 1st January 1979 to BAe with the Hawk, to Lockheed with the Dassault-Dornier Alphajet, and to Advanced Technology Systems, USA, with the Aermacchi MB339, for studies into the conversion of their existing trainer aircraft. In January 1979, BAe sent out a questionnaire to a number of U.S. companies with a view to ascertaining if any of them were interested in a teaming arrangement. The following companies replied, with varying degrees of interest:-

General Dynamics High interest

Rockwell Awaiting developments

DAC (or MDC, McDonnell Douglas Corp.) No interest, would work alone

Ling Tempco Vought Medium interest

Fairchild No interest.

Northrop High degree of interest In July, NADC reported that unless modified for higher lift, the existing aircraft conversions would have approach speeds 4 to 5 knots higher than the maximum specified value of 115 knots. It also concluded that the weight penalty for modifying the aircraft for carrier operations would entail an increase in take-off weight of some 8 to 10 %, but crucially, the converted aircraft would be significantly cheaper than an all-new aircraft, and involve less risk. This carried great weight in the cash-strapped Defence Budget at that time. In December 1979, an RFP for VTXTS exploration studies was issued, and on 6th February, DAC and BAe announced a teaming agreement, despite the former’s earlier stated intention to work alone, with the Sperry Co. to look after the simulators, etc. In March 1980, the BAe response showed that the approach speed was indeed the critical feature, and gave a value of 108 knots TAS based on achieving a maximum lift coefficient (CL MAX) of 2.2, which they believed they could get with modifications to the existing Hawk wing (restoring the full span of the flap vane and changing the wing leading edge dressing – fences, etc.). In fact this turned out to be very difficult to achieve; the value was based on only one aircraft’s results, and it was an exceptionally good one. (This is discussed in a later section.)


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On the 14th July 1980, a ‘Best and Final Offer’ was proposed by BAe, who now amended the approach speed to 113 knots, having recalculated the speed (VPA MIN) governed by a ‘pop-up’ manoeuvre (see later), but still based on achieving a CL MAX of 2.2. On 21st August 1980, six contract study awards were made, 3 for new aircraft and 3 for derivatives; DAC/BAe were awarded one in each category. For the next year DAC, MDC and BAe engineers were involved in extensive discussions on the value of CL MAX and the corresponding VPA MIN for the final submission.

The “Best and Final Offer” was submitted by DAC (as prime Contractor) on 28th September 1981 with a VPA MIN value of 117.9 knots TAS, based on a CL MAX of 2.1, which was DAC’s

re-computed value based on BAe data. On 18th November 1981, DAC (MDC) / BAe were selected (over 5 other candidates) by the US Navy, to supply the VTXTS training system, including an aircraft element of 253 T-45A aircraft, as they were to be known. 6.4 Design Changes The requirement to carry out catapulted take-offs and arrested landings from U.S. Navy aircraft carriers (CARQUALS) resulted in very significant design changes to the basic Hawk, taking the Mk.53 as the starting point. First, and most obviously, was the upgrading and strengthening of the undercarriage. Aircraft are ‘flown into’ the deck of a carrier without any attempt to ‘round-out’, so that a final approach speed of say 120 knots TAS along the specified 4 ˚ descent flight path imposes a possible 14 ft/sec vertical velocity on the main undercarriage. This is much higher than on the land-based aircraft, so a much sturdier and longer stroke main undercarriage was needed. In turn, this required the main undercarriage leg to be mounted further outboard. The necessity for a catapult tow hook on the nose undercarriage, to which the steam catapult strop would be attached, plus a ‘hold-back’ bar, meant a complete re-design of the nose undercarriage. The single nose wheel had to be replaced by a twin steerable unit and extra structure provided to take the tension from the pull of the catapult into the main fuselage structure. Approximately 900 lb of weight was added to the basic aircraft by these changes. The acceleration due to the catapult launch also meant that the aircraft control operating rods had to be balanced in the longitudinal plane; under acceleration the action of some of the original unbalanced rods were reversed. At the rear of the aircraft the over-riding requirement was the provision of an extending arrestor hook mounted underneath the rear fuselage centreline. This meant that the airbrake had to be moved from its position under the fuselage on the Hawk. It was split into elements on each side of the fuselage, just ahead of the tail, operated by twin hydraulic jacks to open to about 60 o deflection. These were mounted external to the fuselage skin and had perforations through them to adjust the aerodynamic effects. The two ventral strakes found necessary on


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the Hawk were replaced by a smaller single ventral strake mounted just in front of the massive hook attachment. The directional flying qualities were enhanced by the provision of an augmentation system for the mechanically powered rudder (the other control surfaces are hydraulically powered). The vertical fin is slightly taller than on the T.Mk.1 to balance the destabilizing effects of the larger main wheel doors when open, allowance being made in the structure for the larger fin loads. The wing was basically that of the Hawk, but structurally modified to cater for the new undercarriage and higher loads. The double-slotted flaps had the full-length flap vane and tailplane canard vanes (‘SMURFS’ – Side Mounted Upper Rear Fuselage Strakes - or ‘stabilator vanes’, in US parlance) were fitted ahead of the tailplane (‘stabilator’). During development the tailplane was extended in span by 4 inches per side, with a corresponding small increase in area and the wing tips were squared off, also giving a small increase in area and slightly higher lift at the cost of transonic performance, which was of less importance in this aircraft. But the biggest change to the wing occurred during later development, when full span leading edge slats were designed by a joint Mc.DD / Brough team, using up-to-date CDF (computer fluid dynamics), backed by tests in the 5 metre low speed wind tunnel at RAE Farnborough on the 1/3rd scale Hawk model. This was a modified version of the joint MoD & BAE project produced earlier by the wind tunnel technicians at Brough. The slats were required to meet high lift requirements and to cater for weight growth – this is discussed later. There were, of course, significant changes to equipment due to Navy demands, chief among them being the provision of Martin-Baker NACES ejector seats having zero height, zero speed, safe ejection capability and an on-board gaseous oxygen system (OBOGS). The engine (F 405-RR-401) was originally based on a Adour Mk.861, but this was upgraded to the Mk.871 of 5,845 lb. uninstalled sea level static thrust (SLST). It included some different control features to enhance its power-up time. The quoted, installed, thrust is given as 5,527 lb SLST on a standard day. The front and rear cockpits are differently equipped from the Hawk, and more recent aircraft have ‘glass cockpits’ with electronic instrumentation. Construction of the new front fuselage, including the cockpits, was the responsibility of MDC (now at St Louis), with the other main components provided from the U.K., comprising about 70% of the work share. Assembly and test of production aircraft was also at St Louis. Figures 46 and 47 illustrate some of these changes.


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Figure 46 T-45A on approach to carrier deck Some of the modifications can be seen – twin nosewheel, longer main legs, arrester hook down,

side mounted airbrakes open, stabilizer vanes. The big modification is the slat on the leading

edge, with a small breaker strip inboard.

Figure 47 This rear view shows how much the oleos collapse on the runway; also the ventral

fin, the retracted hook, air brake perforations and the ‘luggage pod’ 6.5 Development.


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From November 1981 up to the go-ahead for Full Scale Development (FSD) in October 1984, engineers on both sides of the Atlantic worked to resolve issues around the value for VPA MIN and the associated values of CL MAX needed, and how to obtain them with the Hawk wing.

Some details of these problems are discussed in the Appendices given later in Part 2. The first two prototypes (162787 and 162788) were started in February 1986 and flew from Long Beach, California, in 1988, on 16th April and 2nd November respectively, in the hands of DAC test pilots. Funding was granted for the first 3 production lots (12 aircraft) three months later. At the end of 1989 the entire T-45TS work was transferred to St Louis, and in September and October 1990 the first production aircraft made their initial flights. Within a month or two, 2 aircraft were delivered to the Navy Air Test Centre at Patuxent River and on 14th December 1991, the first carrier landing was made on the USS John F. Kennedy. Navy test pilots came up with a number of deficiencies as a result of their flight trials on the prototypes of the T-45A. The five most serious, and the corresponding corrective actions, were as follows:

(i) A longitudinal control instability at high Mach number – this was corrected by changing the stick-tail gearing and the provision of dampers,

(ii) Engine thrust was too low at high atmospheric temperatures, and the idle RPM was too low for the approach. The engine acceleration time to full thrust from approach power was too long. A de-rated (for longer life) Adour 861 had been installed; this was replaced by a version of the Adour 871 (F 405), with a modified fuel system to improve the ‘spool-up’ time.

(iii) There was excessive pitch change on extending the side mounted air brakes – these were subsequently inter-connected with the tailplane. (A similar scheme had been developed and flown on the Hawk T.Mk.1, but not implemented).

(iv) The lateral / directional stability was inadequate. This was corrected by the provision of a large central ventral fin, a taller vertical fin, an aileron-rudder interconnect and a yaw damper for low speeds.

(v) Inadequate stall performance and handling. As already mentioned the wing tips were squared off and, most significantly, full span leading edge slats were fitted. (The cost of providing these was, inter alia, the subject of a court case, which is mentioned later, Section 6.9). A technical evaluation (TECHEVAL) was held towards the end of 1993, and an operational evaluation (OPEVAL) early in 1994. The T-45 came through these very successfully, with the quote “the T-45TS is determined to be operationally effective for ground school familiarisation, basic instruments, radio instruments, airways navigation, instrument rating, formation and night familiarisation training”. 6.6 Production.


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The first production aircraft (163601) from MCAIR, St Louis, flew on 16th December 1991 and was handed over about a month later. Initially, production was planned as one aircraft per month. Production from an assembly line at Palmdale, California proceeded comparatively slowly from 1992 onwards. Production ceased in October 2009, the last aircraft (No.167106) being handed over on 20th October 2009; 223 aircraft had been produced. Procurement of production aircraft by fiscal year (FY) (also known as financial year) is given in the table below.

Fiscal year 1988 1989 1992 1993 1994 1995 1996 1997 1998

Aircraft procured 12 24 12 12 12 12 12 12 15

Fiscal year 1999 2000 2001 2002 2003 2004 2005 2006 2007

Aircraft procured 15 15 14 6 8 14 10 6 10

Total 221 aircraft procured (plus two prototypes) Early aircraft did not have leading edge slats, but these were retro-fitted to all aircraft. Originally, 54 ‘land-based’ aircraft without the slatted wing were planned but this was rescinded, and all aircraft were to be carrier-capable (see also Section 6.9). A digital ‘glass’ cockpit (Cockpit 21) was fitted to the 37th aircraft, which first flew on 19th March 1994. The fitment was finally approved from the 84th aircraft, rolled out at the end of October 1997, and validated in 1998. It was authorised to be retrofitted to earlier aircraft in fiscal year 2003 (FY 03). These aircraft were referred to as the T-45C. Starting from 2000, all T-45A aircraft coming under routine servicing were supposed to be fitted with new, extended upper lips to the intakes. This was to improve the air flow at higher angles of attack. Though extensive calculations and some model tests were done, it is believed that these were never flown. Engine surges experienced under these conditions were dealt with by an engine modification. 6.7 Future intentions. Nineteen T-45C aircraft were to be fitted with a ‘virtual mission training system’ with an intention for initial operations in 2011. A new version, T-45D, had been proposed in discussions between USN, Boeing and BAE Systems, intending to procure 180 aircraft having life extensions to 2040 but this has not been proceeded with. 6.8 Operational History. September 1994 First student class graduates. March 1999 200,000 fleet flying hours reached. April 2000 1,000th student graduated.


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April 2001 First aircraft exceeds 5,000 hrs. May 2001 342,000 hours, 22,000 deck landings, 1,328 students graduated. September 2003 500,000 flight hours. February 2008 800,000 flight hours, 3,000 students graduated. October 2009 932,000 + hours, 3,600 students graduated. 6.9 Litigation. Faced with cost overruns rumoured to be of the order $400 million, the US Navy argued in legal proceedings in the U.S. courts for repayment of the money by the McDonnell Douglas Corporation (MDC) on the basis that the contractor had failed to meet the contractual specification. What is presented in this section is a verbal report on the case, taken mainly from witnesses’ memories. It does not pretend to be a rigorous report on the proceedings. The crux of the issue was that the T-45, in its original form, did not meet the specification in terms of VPA MIN. A maximum value had been specified and the means for determining it had

been laid down. The critical technical feature of this was the achieved value of maximum lift coefficient CL MAX in the carrier approach configuration, and its determination from flight

tests. The Navy contended that the leading edge slats were fitted by MDC in order to meet this deficiency, whereas MDC contended that the requirement could have been met without the slats, though again, this depended on how the stalling speeds were derived, and at what landing weight. Towards the end of 1995, formal statements under oath were taken from the BAe personnel in premises at Warton, Lancashire, and elsewhere. These were submitted to the US court. In October 1996, BAe witnesses were informed that the case between USN and MDC had been resolved, very much to MDC’s satisfaction. The clinching argument seemed to be an instruction given (called ‘a constructive change’) at an MDC / USN meeting that MDC should fit slats on the leading edge of the Goshawk. This was claimed by MDC to be a direction to them by the Navy Configuration Action Team. Accordingly, in the autumn of 1989, slats were designed by MDC and BAe, made, fitted, tested and cleared for use by late summer. The first production wing with a slat arrived at St Louis in mid-1991. The USN was not able to refute this claim and so settled out of court. Nevertheless, at the end of all this, the US Navy procured a highly capable, efficient training aircraft, whose use may well extend into 2040. 6.10 Final Comments. The conversion of a fairly simple land-based trainer to a fully capable carrier-qualified naval aircraft was declared by some to be impossible, and that a new aircraft was necessary. The original design team together with their colleagues at Brough showed that this was not so, and their efforts seem to have been very successful, according to reports of the US operations.


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7. LEADING PARTICULARS AND PERFORMANCE SUMMARY 7.1 Leading Particulars of the T-45A “Goshawk” Wing span 30ft. 10 in. (9.39m) Wing gross area 190 sq.ft. (7.66 sq.m.) Wing sweepback on ¼ chord line 23.7˚ Stabilator (tailplane) span 15ft 1in (4.59m) Stabilator area 47.6 sq.ft. (4.43 sq.m.) Stabilator sweepback, ¼ chord 30.1˚ Vertical stabilizer (fin) area 27.9 sq.ft. (2.59 sq.m.) Vertical stabilizer sweep, ¼ chord 39.5˚ Vertical stabilizer height 6ft. 4in. (1.93 m.) Operation ready weight, 2 crew, no fuel or pylons 10,403 lb (4719kg) Usable fuel capacity 432 US gall (360 IG) (1636 litres) 2,937 lb at 6.8lb/US gallon Normal take-off weight (Clean) 13,340 lb (6050 kg) Maximum take-off weight up to 15,000 lb Weight with 60% fuel (clean) 12,167 lb (5519 kg)

7.2 Field Performance. Stall speeds at 12,000lb gross weight:

Power off Approach power Max power

Flaps up, gear up (CAS) 120 kn 119.5 kn 109 kn

Flaps T.O., gear down 103 kn 101.5 kn 93 kn

Flaps Ldg., gear down 97.5 kn 96 kn 90 kn (Equivalent CLMAX 1.96 2.02 2.27)

Take off in still air at 13,340 lb (estimated). Lift off speed 126 kn CAS. (flaps T.O.)

Take-off distance to 50 ft 3,420 ft 3,900 ft with take-off flap setting at SL, ISA at SL, ISA +15˚C with landing flap setting Approx. 80% of the above values.

Landing (estimated) in still air. 11,000 lb 12,000 lb

Approach Speed 114 kn CAS 119 kn CAS

Ground roll, dry runway, SL, ISA 3,100 ft 3,300 ft

Ground roll, wet runway, SL, ISA 4,800 ft 5,200 ft


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7.3 Range Performance. Note: This is presented as the estimated “No Allowance Still Air Range” (NASAR) This assumes that all the fuel available is used under cruising conditions. When

allowance is made for take-off, climb and descent, the practical range is approximately equal to 75% of the values shown. The take-off weight is taken as 13,340 lb and the total fuel as 2,940 lb. No pylons or stores are fitted.

Optimum cruise conditions. Mach 0.716 (410 kn TAS) Mean Altitude 39,200 ft.

NASAR 1158 nautical miles Cruise at constant altitude, at best speed. Altitude (ft.) Mach no TAS kn. NASAR (nm)

Sea Level 0.35 216 547

5,000 0.38 247 620

10,000 0.42 268 700

15,000 0.46 288 767

20,000 0.51 313 879

25,000 0.56 347 958

30,000 0.61 359 1,053

35,000 0.675 389 1,147 Cruise-climb at constant True Air Speed. Altitude TAS Mach NASAR TAS Mach NASAR (ft) (kn) number nautical miles (kn) number nautical miles

Sea Level 350 0.53 488 500 0.76 319

5,000 ft 350 0.54 563 500 0.77 370

10,000 ft 350 0.55 645 500 0.78 429

15,000 ft 350 0.56 744 500 0.80 498

20,000 ft 350 0.57 851 500 0.81 571

25,000 ft 350 0.58 963 500 0.82 638

30.000 ft 350 0.59 1060 500 0.85 680

35,000 ft 350 0.61 1113

40,000 ft 350 0.61 1027


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7.4 Turning Performance. 12,500 lb, ISA, no pylons or stores 5,000 ft 15,000 ft

Max. sustained turn rate (deg/sec) 12.5 8.7 at Mach number 0.4 0.5

Max. sustained normal acceleration 4.6 g 3.4 g at Mach number 0.72 0.75

Min. turn radius (buffet onset) 2,200 ft 3,200 ft

Lift /buffet limited turns:

Max. turn rate (deg/sec) 16.5 12.5 at Mach number 0.65 0.75

Max. normal acceleration 6.5 g 5.7 g 7.5 Maximum Level Speed. Maximum thrust, ISA day. 13,000 lb gross weight, no pylons or stores. Altitude (ft) Mach number True Air Speed (kn.)

Sea level 0.81 536

10,000 0.84 536

20,000 0.85 522

30,000 0.856 505

40,000 0.83 476 7.6 Structural and flight limits. Max. normal acceleration, low level 6.5 g 10,000 ft and above 7.33 g Max. negative normal acceleration - 4 g Max. speed 550 kn CAS Max. Mach number 1.04


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APPENDIX 2.1 THE MINIMUM POWERED APPROACH SPEED VPA MIN A2.1.1 BACKGROUND. The derivation and determination of the minimum speed permissible in the approach to a carrier landing was perhaps the single most demanding requirement of those needing to be met in the T-45 Specifications. Many man-hours, on both sides of the Atlantic went into the solution of the issues involved and their deliberations had a profound impact on the design and eventual construction of the aircraft. In attempting to make a successful landing on a comparatively small carrier deck, in the midst of a wide and featureless ocean, the naval aviator is faced with a difficult precision task, and every effort must be made in the design of a carrier-capable aircraft to help him. He has to maintain a steady flight path of descent to the carrier deck and to pick up one of several arresting wires stretched across that deck. Various techniques are employed to assist him to do this, but probably one of the most important criteria is the speed at which he approaches, and hence flies on to the deck. It is important to keep this speed to a minimum, not only to give the pilot time to assess and if necessary correct the situation, but to keep the arresting loads as low as possible. To this end, the U.S. Navy laid down the definition of the maximum allowable speed in a powered approach, labelled VPA MIN.

This definition is given below, and is closely related to the maximum lift that is available from the aircraft at low speed in the landing configuration. This is discussed in Appendix 2.2. A2.1.2 Design Minimum Speed in a Powered Approach. This is defined as 1.05 x VPA MIN in the landing configuration, flaps and gear down and power

to maintain a 4 degree downward flight path, the glide slope. The atmospheric conditions are set as sea level pressure and 90˚F air temperature, and the specified speed is given as True Air Speed (TAS), which under those conditions is about 3 % higher than the CAS, the corresponding speed in the Standard Atmosphere. VPA MIN was defined as the highest of the airspeeds defined by the following criteria:

(i) 1.1 x VSPA where VSPA is the power-on stalling speed using the power required for level flight at 1.15 x VSL where this is the power-off stalling speed in the landing

configuration.

(ii) The lowest speed at which the aircraft is capable of making a glide slope correction from one stabilised at VPA MIN to another parallel glide slope 50ft above it within five

seconds of the initiation of the manoeuvre. No change of thrust is permitted during the manoeuvre, and it must be accomplished using no more than half the maximum increment of lift available at initiation. The manoeuvre is considered complete when the aircraft has intercepted the higher glide slope, and it must be capable of maintaining this new glide slope, with the pilot permitted to change the thrust setting as required.


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Clearly, (ii) is very much related to the optimum usage of the longitudinal control, in this case the tailplane, or ‘stabiliser’ in U.S. parlance, and it was shown theoretically, using simplifying assumptions, that the ratio of steady approach speed to stall speed was sensitive to loss of speed in the manoeuvre, having values of between 1.12 and 1.17 for reasonable assumed values of retardation. But this was much too imprecise, and efforts were made to mathematically model the aircraft and predict its flight path in a response to stabiliser movement of the required amount to apply the specified normal acceleration increment. This, in itself, took considerable time and effort, and was the subject of much discussion between U.S. Navy and MDC and BAe engineers, with final understanding reached much later. A2.1.3 Preliminary flight tests. It was recognised early in the process that the establishment of VPA MIN was going to be critical,

and some flight tests were specially mounted to try to shed some light on this measurement. These were conducted on the first RAF Hawk XX154, modified to a configuration giving higher lift – full span flap vane and a standard fence with a small breaker strip adjacent to it. The values of maximum lift coefficient that were obtained were estimated as between 2.0 and 2.2 power off (but the definition of stalling speed was questioned, see Appendix 2.2 for a more detailed discussion on this topic). The tests were carried out in February 1979 at the airfield at the Royal Aircraft Establishment, Bedford, where the long runway was equipped with kine-theodolite cameras. These were able to give, very shortly after a flight, details of the speeds and positions of the aircraft as it carried out a glide slope correction manoeuvre. The test results were reported by S. M. Gerrard, the flight observer involved, in Report KFT-N-Haw-00089 dated June 1979. Eight filmed test runs were made on Flight 406, but the weather was poor with gusts of up to 20 knots, and not all the results could be used. It was concluded that for masses of 4,600 kg (10,140 lb) and 4,800 kg (10,580 lb) the values of 1.05 VPA MIN were 107.2 kn IAS and 109.6 kn IAS respectively, or when related to the specified

90˚F atmospheric temperature, 110.3 and 112.8 kn TAS. (When corrected to the value of landing mass finally used, these become 118 kn TAS). Unfortunately (see Appendix 2.2) it was found that XX 154 gave exceptionally high values for CL MAX and these could not be

accepted for standard production aircraft, so the results obtained had to be discarded. Nevertheless, they showed that provided a value of CL MAX of about 2.0 could be obtained,

then the specified minimum approach speed criterion could be met.


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APPENDIX 2.2 THE ANALYSIS AND DETERMINATION OF MAXIMUM LIFT COEFFICIENT

A2.2.1 The definition of stalling speed. For the Hawk T.Mk1 for the RAF, the stalling speed was defined as ‘the minimum speed reached in a stall manoeuvre with the speed decaying at a rate of 1 knot per second’. Various power settings could be used. This definition was in line with the definition used by international Civil Regulatory Authorities (including the Federal Aviation Agency in the USA) as being easily repeatable and clear. Take-off and landing speeds were related to it, the margins being fixed from considerations of probability of accidents. In particular, the landing approach speed was fixed as 1.3 x (stalling speed in the landing configuration). Note, however, that this did not give the ‘aerodynamic’ stalling speed and its corresponding value of maximum lift coefficient, CL MAX, because usually the normal acceleration at the stall,

as so defined, could be less than unity and any thrust vector was not excluded. Furthermore, to obtain the true ‘aerodynamic’ value of CL MAX, maybe to compare this with a value from a

wind tunnel test, the rate had to be much slower, perhaps half the value used. Thus the value of maximum lift coefficient came out as a lower value than that defined by the minimum speed in a stall manoeuvre, by as much as 10%. The US Navy was justified in using the criterion giving higher stalling speeds because the stipulated approach speeds were also a lower multiple of these, typically 1.15 x stalling speed. This was particularly important when the approach to the carrier case was considered, as the entry speed had to be kept as low as possible, typically only 1.15 Vstall , compared with the

civil 1.3 multiplier (but of a lower stall speed). In the case of the Hawk, the two definitions resulted in a difference of about 0.1 in CL MAX or about 3 knots at the stall.

As discussed in Appendix 2.1, the value of the approach speed was directly related to the stall speed as defined above. A2.2.2 Flight test values of CL MAX for various Hawk aircraft. It is shown in Appendix 2.1 that to meet the specified approach speed a ‘real’ CL MAX of about

2 was needed, relating to about 2.1 using the ‘minimum speed’ definition. This in turn depended on the agreed value for the appropriate aircraft weight in the approach with 60 % fuel remaining, and it was felt that a figure of 1.9 might be used to offer guaranteed values of approach speed, to be on the safe side. Thus flight tests were performed on a number of different Hawk aircraft with different wing ‘dressings’, seeking to confirm that at least a value of 2.0 for CL MAX could be justified. In all

cases the handling at the stall had to be acceptable, with no excessive roll-offs and adequate buffet warning of the approach to the stall. The table below gives some details of the flights carried out and the results obtained. All tests were conducted in a similar way, and all were in the landing configuration.


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Aircraft Flap Leading edge

devices Flight No.

Handling Buffet margin knots

CL MAX

(Min Speed)

XX 154 FSFV Small b.s 416 Good 8 2.32

+fence Small b.s 417 Acceptable 5 2.18

CBFV Small b.s 425 Acceptable 7 1.99

G-HAWK FSFV Fence 587 Acceptable 3 - 5 1.845

623 Acceptable 3 - 4 1.855

632 Acceptable 3 - 4 1.992

633 Acceptable 2 - 4 1.993

Turbs 634 Good 4 2.077

Fence and turbs 637 Good 3 - 4 2.125

b.s. added 638 Acceptable 5 2.12

Various arrays turbs 640 to

661 Acceptable to

good 2 to 5 1.95 to 2.2

average 2.06

CBFV RAF 667 Good 3 - 4 1.688

692 Good 3-4 1.685

XX 154 CBFV RAF 450 Acceptable 6 - 7 1.697

3 mini fences and

outer fence 455 Good 2 - 4 1.886

No outer fence 457 Good 3 - 4 1.90

CBFV 3 mini fences 785 Good 1 - 2 1.80

Repeat 793 Acceptable 1 - 3 1.85

Repeat of flight 455 813 Good 3 1.67

XX 154 CBFV RAF 450 Poor 6 - 7 1.7

G-HAWK RAF 453 Acceptable 4 - 6 1.61

XX 338 (Production)

RAF 16 Good 2 - 4 1.57

Notes: XX 154 seemed to be particularly good. XX 338 was poor and needed rectification.

Flaps CBFV Cut back flap vane; FSFV Full span flap vane

Leading edge devices Turbs Button type turbulators placed round leading edge

b.s. Breaker strip

The penalty in CL MAX for the RAF configuration was about 0.2 to 0.3 (up to about 5 kt.)

In most cases buffet warning was low, but acceptable due to the docility of the stall Despite the very repeatable technique in the stall manoeuvre, there was a degree of scatter in the results which made it difficult to come to firm conclusions.


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A2.2.3 Stall speeds on the Hawk T.Mk.1 and T-45. For the RAF Hawk T.Mk.1, published data quote the flaps down stalling speed as 96 kt, weight unspecified, but possibly for landing with two pilots and typical landing fuel. This implies a CL MAX of 1.6 and an approach speed of 125 kt with a landing speed of about 110 kt at that

weight. In 1976 the author collected together all available results on stalling on instrumented aircraft in the RAF configuration. The results exhibited a fair degree of scatter, but when corrected to a common mass of 4,400 kg, gave the following: VSTALL kt IAS CL MAX indicated

Flaps up, gear up 110 to 116 1.32 to 1.18

Flaps mid, gear down 99½ to 105½ 1.61 to 1.43

Flaps full, gear down 94 to 98 1.83 to 1.69 Since the position error of the instrumented pitot / static was small (within 1 kt), IAS may be taken as CAS without much error. Comparing these figures with appropriate ones in the table above for the RAF configuration indicates that perhaps the upper line through the scatter band of speeds should be used. During the test series, it was found that the position and shape of the breaker strips on the leading edge was quite important, both for stall speed and behaviour. Even a 1 mm reduction in their heights, or any rounding of their sharp leading edges, gave erroneous results. For the T-45, CL MAX with approach power is estimated from given stall speeds as 2.0, but this

was with leading edge slats extended. Wind tunnel tests indicated that the slats supplied an increment of about 0.1. But it was known that a wing with a fixed, drooped leading edge, as flown on the 100 and 200 series Hawks, gave a considerable improvement of maximum lift, so that MDC were indeed justified in claiming that an acceptably high maximum lift could be reached without the fitment of slats.


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APPENDIX 2.3 TAILPLANE STALL ON THE T-45 A2.3.1 Checks on tailplane stall phenomena. The occurrence of tailplane stall is well covered in Part 1, Appendix 1.3, with regard to the later versions of the Hawk, and the function of the ‘tailplane canard vanes’ is explained. There was some concern that, on the T-45, the side-mounted airbrakes might have an adverse effect on their efficiency. Thus two sets of wind tunnel tests were proposed, one set using the high Reynolds Number (R.N.) facility at the Royal Aircraft Establishment, Farnborough, where the Mach number and Reynolds number matched the full scale aircraft at the stall, and another set using the 13ft x 9ft low speed wind tunnel originally located at Weybridge, having only half the Reynolds number capability. Both facilities would use the 0.3 scale model which had been jointly funded by MoD and BAe. The former tunnel had already been used to build up a data set for the T-45, and it was of first importance that the other wind tunnel, more accessible and cheaper to operate, was shown to give comparable results before the effect of the air brakes was assessed. It was found that the curves of lift and pitching moment were almost identical over the working range of +8 to -8 degrees of angle of attack. As was expected, the lower Reynolds number tunnel gave a lower maximum lift coefficient, since the flow separated at a lower angle of attack. Putting in an experimental screen to increase the turbulence in effect increased the Reynolds number in the Weybridge tunnel. Extensive testing then showed that the operation of the airbrakes had very little effect on the action of the TCV – there was no potential problem.


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PART 3 THE HAWK - CURRENT AND FUTURE PROSPECTS SUMMARY This part relies heavily on the information given by Professor Andrew Bradley, Chief Engineer, Hawk, in his presentation to the Royal Aeronautical Society Historical Group on 20th October 2011, supplemented by published data, mainly from various editions of Jane’s ‘All the Worlds Aircraft’ and from BAE Systems. It outlines the development of the Hawk over about the past fifteen years, and discusses the possible future sales for the aircraft in what is now a global industry. 8.1 Events from the mid-90’s to 2011 At the beginning of this period, in October 1992, the Kingston site was sold off to become a housing estate, with the design staff moved to new premises in the Farnborough Aerospace Centre, Hampshire. All technical work except Flight Test and most manufacturing had moved to Brough in East Yorkshire, with the Warton unit in Lancashire carrying out final assembly and flight testing. Later, final assembly and first flights (delivery to Warton for subsequent production flight acceptance) were done at Brough. The last four of 24 UK-built Mk.132 and all but two of the RAF T.Mk.2 aircraft had their first flights at Brough. The aircraft was taken over by new people, though a few ‘old stagers’ came with it. Thus, fresh insights and enthusiasm for the aircraft were brought to bear and there is no doubt that a new impetus was added. The task was to ensure the steady development of the aircraft for the available markets, to keep the aircraft up to date and saleable. The aircraft had to evolve to meet the needs of new customers by bringing in new technology as it became available. It was recognised that the industry was going global, and that the customer, in many regions, wished to produce the aircraft in their own country, in order to develop their own industry. The main vehicles involved were the Hawk 100 series, as a dual role trainer / combat aircraft, and the single seat Hawk 200 series as a light combat aircraft. Based on the initial designs from Kingston as outlined in Part 1, further development continued on these types, as described in the following sections. 8.2 Development of the Hawk T.Mk.1 in RAF service With the steady usage of flying hours by the fleet, a re-wing programme was initiated in 1988, using as a basis the then current wing structure for export aircraft. The first batch of new wings was fitted to 85 aircraft by the end of 1993. A second batch of 59 wings was supplied by 1995. In addition to this, 80 aircraft were rebuilt by replacing the centre and rear fuselages with standard Mk.60 series components. These modifications extended the safe life of the structure, the former 6,000 hours being extended to 10,000+ hrs. These numbers included 11 T.Mk.1’s, 62 Mk.1A’s and 7 Mk.1W (re-winged).


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XX 348, the first aircraft so treated, was delivered on 11th April 2000, and the last of the batch, XX 242, was delivered to the Red Arrows on 27th August 2003. In 2006, the UK MoD fleet numbered 132 aircraft, and on 5th July 2006, a total of 1 million flying hours was achieved. 8.3 Export aircraft In 1992, the company Series 100 Hawk ZJ 100 flew with a new avionics suite, an updated engine and a new wing having a cambered nose section and wing tip missile stations, and this was followed by the single seat demonstrator with the radar equipment in the nose. Between 1993 and 1995, four Mk.103 and 12 Mk.203 (single seat) aircraft were delivered to Oman, 18 Mk.102 to Abu Dhabi and in 1994-1995, 10 Mk.108 and 18 Mk.208 aircraft were delivered to Malaysia. At about the same time, Indonesia, having received 20 Hawk Mk.53 in the early 80’s, followed up with an order for eight Mk.109 and 32 Mk.209 aircraft, all delivered by 1996. In 1996, the Company was authorised to initiate development work on a type of Hawk for the Royal Australian Air Force (RAAF). In 1997 an order was placed by the RAAF for 33 Mk.127 aircraft (Figure 48). These were to be assembled in a BAE SYSTEMS factory located in Williamstown, NSW.

Figure 3.10 Hawk Mk.127 Royal Australian Air Force Livery for 60th Anniversary of 79 Sdn. Note wing tip CATM (Captive Air training Missiles) and 130 IG external tanks.


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Also in 1997, ZJ 100 was fitted with a ‘Glass cockpit’, and the Mk.127 had a 3MFD ‘glass cockpit’ together with On Board Oxygen Generating System and AC generation. It retained the Adour 871. The Mark 129 is similar, but with the Adour 951. Hawk 100 series orders continued with a contract in 1998 from Canada for 22 Mk.115 aircraft, basically a Mk.109 with Adour 871 engine and DC generation, and in 2000, another contract from South Africa for 24 Mk.120 aircraft (similar to the Mk.127 but with a Adour 951), which were to be assembled by Denel Aviation, Johannesburg. The Mk.120 had a mission system from a South African source. In 2001 the company started to investigate a new RAF training system – see Section 8.8. Bahrain ordered six Mk.129 aircraft in 2003 and in 2004-2005, an order was received from India for 66 Hawk Mk.132. Twenty-four of these were built in the UK, six assembled by Hindustan Aircraft Ltd., Bangalore, from kits, and 36 built by HAL This was followed in 2010 by a further order for 57 aircraft, all built under licence by HAL. Many changes had been made to the aircraft, engine and systems – these are outlined in the next Section. 8.4 The Aerodynamic Development of the 100 and 200 series aircraft . Responsibility for Hawk Design passed to Brough in 1986-88. As related earlier, some development work had been carried out at Kingston and Dunsfold on series 100 and 200 prototypes and this work was now taken over at Brough and Warton. The aircraft used were ZA101, ZJ100, ZG200, ZH200 and ZJ201. The most noticeable external change was the provision of wing tip stations for missiles and a longer (Laser/FLIR) nose. The original leading edge fixed droop was re-designed and increased to improve high speed behaviour, and enhanced turning performance was obtained by the addition of a combat flap setting (nominally ¼ flap) usable up to 350 knots IAS / 0.8 Mach number. A minor transonic pitch-up was cured by a modification to the wing vortex generator arrangement. Whilst the modified wing tip did not produce adverse aerodynamic effects, it was found necessary to alter the wing fence arrangement to improve stall handling. These may be summarised with reference to the Hawk T.Mk.1, as below:- T.Mk.1 (cut back flap vane, rounded wing tip)

Large fence at 3,165 mm, two 98 mm breaker strips on leading edge, inner edges at 1,500 mm and 2,535 mm.

Early series 100 (longer nose, full span flap vane and SMURFs, drooped LE. Wing tip missile stations)

Three mini fences at 1,500 mm, 1,880 mm and 2,220 mm, plus outboard mini fence at 3,507mm


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Mk.109 series onward.

Large fence at 2,505 mm, small span L.E. breaker at 1,600 mm, mini fence at 3,807 mm With its longer nose and redistribution of equipment, the inertia changes of the aircraft were such as to cause oscillatory spins to be produced. As the requirement was for the aircraft to have robust and repeatable spinning behaviour, wind tunnel testing helped to develop a new and carefully defined fence position. This was successfully flight tested. A powered rudder/yaw damper system was introduced and this gave rise to the unexpected occurrence of rudder “buzz” in dives to Mach number 1.2 on the single seat aircraft. This was found to be due to aerodynamic effects on the flow round the rudder hinge, coupling with the frequency of the powered rudder system. A change to the ILS aerial at the top of the fin cured the problem. Another problem on the single seat aircraft was to develop a new pitot / static system – the single element used on the two-seaters could not be used as the single seat aircraft had a radar nose. A multiple unit system was fitted. Much of the aerodynamic development described above was led by the team of aerodynamicists at Brough, with the support of similar staff at Warton. The 1/3 scale low speed model, a joint venture between BAe and MoD originating back in the 80’s, was modified and was tested in the 5-metre pressurised wind tunnel at RAE Farnborough. 8.5 The Hawk AJT (Advanced Jet Trainer). 8.5.1 Description of the airframe. The Hawk AJT is the current version offered to customers, and is billed as the ideal trainer for potential pilots of advanced fast jet military systems. As a combat aircraft in its own right it offers a significant light combat capability. Of the seven store stations, the wing tip stations are for missiles, but the others can be used for a wide variety of modern weapons. (This was envisioned at the earliest stages of the Hawk design at Kingston). Though externally similar to the basic Hawk, there were many internal changes to the structure and equipment in the period under consideration. Most noticeable externally were the increased length of nose to accommodate a new avionics suite, and the installation of FLIR therein. The wing with its seven weapon stations (load up to 3,000 kg) has a fixed drooped nose to improve lift at high speeds, and a different array of vortex generators along the 25 % chord line. There is a small breaker strip inboard of a fence positioned at about mid semi-span. A single ‘mini-fence’, similar in size to one of the three used outboard on the Series 60 wings, is in an outboard position (see above). The full span vane is retained, and the flaps now have four positions, with a ¾ setting and a combat setting which may be used up to a speed of 350 knots. The tailplane canard vanes are standard. A hydraulic yaw damper is applied to the powered rudder.


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The whole airframe structure has been revised to improve ease of production and maintenance; the safe life is extended to around 12,000 hours. A HUMS system monitors component life and engine usage and equipment health. Nose wheel steering has been incorporated on a strengthened nose leg. Aircraft systems are updated, with Martin Baker Mk.10 lightweight ejector seats, on-board oxygen generation (OBOGS), and, to cope with the extra electrical load, an engine-driven 25 kva A.C. generator has replaced the former 9kw D.C. generator. The completely new avionics suite includes an integrated advanced navigation/attack system with two modern mission computers, inertial navigation and ground position systems, radar warning receiver, and chaff and flare dispensers. A plug-in flight refuelling probe may be fitted, and external fuel, generally 2 x 130 imperial gallon capacity tanks on the wing pylons can be supplemented by a 100 imperial gallon centreline tank/baggage pod. Pilot selection of services is improved with a HOTAS (Hands On Throttle and Stick) unit and pilot operated controls on the control column (Figure 49).

Control Column Handle Throttle Handle

Figure 49 Hawk Advanced Jet Trainer

BAE SYSTEMS

The above capability allows the combat version of the Hawk to perform effective combat missions at a fraction of the cost of front-line combat aircraft having one or more re-heated engines.


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8.5.2 Leading Particulars. Wing span 32 ft 7 in (9.94 m) with missiles

Wing area 179.64 sq.ft (16.7 sq.m)

Wing sweepback 26 o at L.E., 21.5 o at ¼ chord

Basic mass empty 10,075 lb (4570 kg)

Design take-off mass 20,062 lb (9100 kg)

Design landing mass 13,007 lb (5900 kg) Normal acceleration limits.

With full internal fuel, no stores +8 to -4 g

60 % fuel; 3,000 lb (1,360 kg) war load +8 to -4 g

60 % fuel; 6,000 lb (2,720 kg) war load +6 to -3 g 8.5.3 Performance. The good handling and performance of earlier Hawks is enhanced in the AJT. The maximum level CAS is set at 560 knots, 0.85 Mach number, with a dive capability to 1.2 Mach number. Take-off performance at Sea Level Clean Aircraft Combat Role 2 crew 1 crew

Full internal fuel Full internal + external fuel ½ flap ¾ flap

Weight 13,292 lb (6,027 kg) 17,793 lb (8,071 kg)

ISA (15˚C) Ground Roll 1,919 ft (585 m) 3,494 ft (1,065 m) Distance to 50 ft (15 m) 2,936 ft (895 m) 4,823 ft (1,470 m)

ISA + 20˚ C (35˚C) Ground Roll 2,329 ft (710 m) 4,281 ft (1,305 m)

Distance to 50 ft (15 m) 3,445 ft (1,050 m) 5,774 ft (1,760 m)

ISA + 35˚C (50˚C) Ground Roll 2,838 ft (865 m) 5,298 ft (1,615 m) Distance to 50 ft (15 m) 4,052 ft (1,235 m) 7,070 ft (2,155 m)


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Landing performance at Sea Level Still air, dry runway, brake parachute deployed

Clean aircraft with 2 crew Combat Role, 1 crew

Full fuel 10% fuel 10% fuel, stores gone Weight 13,290 lb 10,754 lb 12,209 lb (6,027 kg) (4,878 kg) (5,538 kg) ISA (15˚C)

Ground Roll 2,871 ft (875 m) 2,100 ft (640 m) 2,493 ft (760 m) Distance from 50 ft (15 m) 6,512 ft (1,985 m) 3,625 ft (1,105 m) 4,101 ft (1,250 m)

ISA + 20˚ C (35˚C) Ground Roll 3,068 ft (935 m) 2,231 ft (680 m) 2,641 ft (805 m)

Distance from 50 ft (15 m) 4,544 ft (1,385 m) 3,625 ft (1,105 m) 4,101 ft (1,250 m)

ISA + 35˚C (50˚C) Ground Roll 3,215 ft (980 m) 2,329 ft (710 m) 2,707 ft (825 m)

Distance from 50 ft (15 m) 4,987 ft (1,520 m) 3,920 ft (1,195 m) 4,413 ft (1,345 m) Optimum Fuel Flow Clean aircraft, 12125 lb 5500 kg, ISA

Sea level at M = 0.3, 200 kn TAS 18.7 lb/min (8.5 kg/min) 10,000 ft (3,048 m) at M = 0.35, 223 kn TAS 16.8 lb/min (7.6 kg/min) 20,000 ft (6,096 m) at M = 0.4, 246 kn TAS 15.7 lb/min (7.1 kg/min) 30,000 ft (9,144 m) at M = 0.5, 295 kn TAS 15.2 lb/min (6.9 kg/min) Sustained Level Turns at Sea Level at optimum speed. Clean Aircraft Combat Aircraft Weight 12,160 lb (5,516 kg) 13,618 lb (6,177 kg)

Without combat flap ISA conditions 6.0 g 4.1 g ISA + 20˚C 5.3 g 3.6 g With combat flap

ISA 15.8 deg/sec 12.7 deg/sec ISA + 20˚C 14.2 deg/sec 11.6 deg/sec Typical Missions. Typical tactical training missions take about one hour with sensor and weapons simulation or with live training stores. As examples of combat operations, an aircraft equipped with a centreline gun and ammunition, and four air-air missiles and 2 x 130 imperial gallon combat tanks can execute a Combat Air Patrol sortie with over an hour on station, 200 miles from base.


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A typical Hi-Lo-Hi sortie carrying 2 x Mk 82 bombs, gun and 2 air-air missiles yields a radius of action of 380 to 440 nm (700 to 800 km) with combat tanks. For Close Air Support, carrying 4 x Mk.82 bombs, 2 air–air missiles and a centreline gun, the radius of action can be nearly 150 nm (300 km), and the Hawk’s low fuel consumption gives it the ability to loiter for long periods, reacting as required to changing battlefield scenarios. With full internal fuel and external fuel in 2 x 130 imperial gallon tanks, the ferry range is over 1400nm (2,600 km), which may be extended further if the centreline fuel tank/baggage pod is used. For the longest ranges, the optional refuelling probe may be fitted for air-air refuelling. Experience has shown that pilots find it easy to use. 8.6 Development of the Adour Engine. The first 100 series NDA Hawk (ZJ 951) flew on 5th August 2002 with an interim standard of a new version of the Adour, the Mk.951. The engine incorporated new technology read across from the Rolls-Royce EJ 200 and Trent engines, and, though it had greater thrust than earlier designs, it also had improved reliability and longer time between overhaul. Full authority digital engine control (FADEC) and an automatic surge recovery systems were provided. The production Mk.951 engine came on line in May 2003. The development of the engine may be seen in the table below: Engine type no. Mk.151 Mk.871 Mk.951 Hawk type. T.Mk.1 Mk.60, 100, Mk.120 + & 200 series series

In-service date 1976 1988 2002 Sea Level Static Thrust (lbf) 5,200 6,030 6,500

By-Pass Ratio 0.79:1 0.76:1 0.78:1

Compressor pressure ratio 10.7 11.3 12.2

Air Mass Flow (lb/sec) 94 97.6 105

Specific Fuel Consumption 0.71 0.78 0.78 (lb/hr/lbf thrust)

Time between overhauls (hr) 1,200 2,000 4,000 8.7 The M.o.D. requirements for a new RAF pilot training system. At the turn of the century, the requirements for future advanced jet aircraft training systems were being addressed by the Ministry of Defence. When formulated, they called not only for the traditional standards of airmanship, but hugely enhanced situational and tactical awareness,


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with training in the management of a digital cockpit, sensors and smart weapons, leading to improved decision making and leadership qualities. It was intended to deliver release to service by August 2008. With competition from the Finmecanica Aermacchi 346 and the LM-KT 50, a £200 million UK product development investment was made, and this has developed the Hawk T.Mk.2 into a market leading standard in sensor and weapons simulaton, datalink networking, mission data recording and de-briefing. 8.8 The Hawk T.Mk.2 (Mk.128) The development of an advanced jet trainer for the RAF, based on the Hawk, began in 2001 with the development of an open architecture mission computer. The Hawk T.Mk.2 is a version of the Lead-In Fighter Trainer (LIFT) and was initially called the Mk.128. It was planned that 31 aircraft with the Adour 951 engine would enter service in 2008. The contract was possibly worth £800 million in total. It is similar to the Mk.127 but with sensor and weapon simulation, autopilot and extra air data probes (Smart Probes) for reduced vertical separation minima compliance, TCAS II (Traffic alert and Collision Avoidance System) with upper and lower antennae, nose mounted conspicuity light and the Adour Mk.951 engine. This Advanced Jet Trainer (AJT – see above) is the standard version now being offered for export. On 22nd December 2004 a design and development contract was let for the provision of two aircraft. The first of these, ZK 010, made its first flight on 27th July 2005, followed by the second, ZK 011, on 6th June 2006. On 19th October 2006 a production contract was placed for a further 26 aircraft. The first production aircraft (ZK 012) made its first flight on 4th August 2008; the release to service was signed on time (to the day) on 30th August. The first aircraft deployed to RAF Valley was ZK 014, on 8th April 2009, and the new flying training system began in late 2009. By 1st January 2010, 23 Hawk T.Mk.2 had flown and 17 had been delivered (Figure 50). The first software update to the T.Mk.2 was released to service in 2010 called ‘Operational Capability 2’ (OC 2) software, the development of which went hand-in-hand with the new RAF training syllabus. The system featured, for the first time, sensor simulation in use in an RAF jet training aircraft. Simulation includes the use of AI radar, radar warning receiver, deployment of chaff and flare, guns, bombs, SRAAM, MRAAM, and ground threats. The whole mission can be de-briefed after flight and analysed by student and instructor. Since it is software based, the system should be reliable, not having any expensive sensor hardware to fail or be maintained. OC 2 has been enthusiastically received. “The Hawk T.2 has revolutionised UK Fast Jet Training and this makes it the optimum Fast Jet training platform” (Group Captain Bruce Hedley, Station Commander, RAF Valley, September 2010).


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Figure 50 Hawk T.MK.2 (Mk.128) ZK 029 Note longer nose, wing tip CATM, centreline tank with baggage compartment.

“It is genius. The new Hawk was streets ahead of its predecessor even before the software upgrade but with OC 2 the pilots are now able to train almost exactly as they do on the front line. They could not be happier” (Brian Braid, OC, 19 Squadron). 8.9 Concluding remarks. The development of the Hawk has been a highly successful and profitable programme for the UK, leading to very significant exports around the world. Currently, efforts are being mounted to promote new buy aircraft in 2012 - 2015. One the most important of these is the T-X requirement of the US Department of Defence for the USAF with a potential order of some 300+ aircraft and a new training system. This is to replace the current training scheme in which the T-38 has served for many years. BAE SYSTEMS has announced that Northrop Grumman will be the manufacturing lead in the joint team and recently it became known that L-3 Link Simulation and Training was joining them, as the provider of the ground based training system for the USAF T-X proposals. In a paper published by the Institution of Mechanical Engineers in 1992, G.Chisnall, of British Aerospace (Military Aircraft) Ltd, North Humberside, remarked: “BAe owes a considerable debt to the personnel involved from both HSA, the RAF and MoD who had the foresight to ensure that the Hawk would not simply be tailored to one specific role but would have development potential built in from the beginning. The current success of the Hawk shows them to have been absolutely correct”. The current team are continuing the tradition with great success.


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Today’s aircraft (2011) still has a family resemblance externally to the original Hawk, and it carries the same ‘genes’ of good handling, brisk and agile performance, low fuel burn, high reliability with low maintenance and excellent visibility from both cockpit positions. For combat, the load carrying ability is impressive, and this was the objective from the beginning of the project. Internally the aircraft is very much improved and the use of sensors and simulation gives a great improvement in versatility in training. With all these changes the weight has, understandably, increased, but so has the thrust of the Adour engine, now with greater TBO (Time Between Overhaul) than before. With about a thousand aircraft ordered or delivered to date to customers worldwide, the aircraft and its training system are a sound and mature basis for further development, as exemplified with the Hawk AJT. With the exploitation of emerging new digital opportunities, there seems to be good career path ahead for the new Hawk. ACKNOWLEDGEMENTS The author is indebted to a number of past colleagues for their recall of events, and particularly to Chris Farara, the curator, for access to the vast store of information on the Hawk in the archives of the Brooklands Museum, Weybridge. Chris Hodson, son of Gordon Hodson also helped considerably as did David Hazzard with photographs. Brian Riddle at the National Aerospace Library, Farnborough, supplied valuable data on aircraft and engines. The figures and photographs have been obtained from various sources, but mainly from material originally published by Hawker Siddeley or British Aerospace, now BAE Systems, and the author is grateful to BAE Systems for permission to use them in this publication. The author thanks all of the above for their help, and must point out that any opinions expressed are entirely his own – nobody else is to blame for any errors!


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As part of the reviewing of this paper, a copy was sent to Duncan Simpson OBE, CEng, FIMechE, FRAeS. He sent the Editor the following letter.


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Paper on the Hawk aircraft for the Journal of the Aeronautical Society

I have read Harry Fraser-Mitchell’s splendid history of the Hawk from initial selection to the present day. Not only is this paper a tribute to Harry, but to the Kingston and ex-Folland design, manufacturing and development team in the early days, [for] taking it through Boscombe Down and entering RAF service in two years two months from first flight. Inevitably it brings back memories of the first aeroplane XX154 arriving at Dunsfold in July 1974 and being worked on, night and day, to ready it for the Farnborough Show at the beginning of September. In fact the first flight took place at 7.21pm on 21st August, and the second – in good weather – the following day. The aircraft flew to Farnborough on its tenth flight and took part in the Flying Display each day – at times in severe weather. It returned to Dunsfold to complete 25 flights before being grounded for two months for comprehensive instrumentation. There had been no unserviceability – a good omen for the RAF fast jet trainer. Two pilots, Andy Jones and Jim Hawkins, were recruited from Boscombe, both Flying and Weapons instructors and they played a major part in obtaining a C.A. Release for the T. Mk 1 within two years. Hawk XX163 was delivered to RAF Valley on 4th November 1976. The Commander-in-Chief [ACM Sir Rex Roe GCB, AFC] had requested to participate in this flight – which he duly did.


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We arrived at Valley, in heavy rain, to be welcomed by Group Captain Thornton and his instructors. So began the illustrious career in the Royal Air Force and overseas services, not forgetting the Red Arrows, of this splendid aeroplane. In T2 form we shall see more of it in the future. The Editor is most grateful to Duncan Simpson for this personal contribution to the history of the Hawk. It brings the story to life and adds an additional view to the history of the technical development of the aircraft told by Harry Fraser-Mitchell.

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The HAWK Story - Royal Aeronautical Society · PDF fileThe HAWK Story Harry Fraser ... apparent that the aircraft then used for the Royal Air Force’s fast jet pilot training were