The HAWK Story - Royal Aeronautical Society

Transcription

Journal of Aeronautical History
Paper No. 2013/01
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 RRTM 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
Paper No. 2013/01
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
1
2
3
4
5
6
7
8
1182-1
1182-2
1182-4
1182-7
1182-8
1182 1182 1182 -
Unswept, high
Unswept, low
Unswept, mid
Swept, high
Swept, low
Swept, low
Swept, low
Swept, high
Tandem
Tandem
Tandem
Tandem
Side / Side
Tandem
Side / Side
Side / Side
Wing root
High
Wing root
Low
High
Pods
Pods
Wing root
U/C
Mounting
Fuselage
Wing
Wing
Fuselage
Wing
Wing
Wing
Fuselage
Notes: All had fixed tailplane with elevators and balanced manual controls.
Unswept wings were rejected as being unlikely to exceed M = 0.8
4
Powerplant
1 x Adour
1 x Adour
1 x Adour
1 x Adour
1 x Adour
2 x BS 153/ Larzac
2 x BS 153/ Larzac
2 x BS 153/ Larzac
Paper No. 2013/01
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
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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
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
0.62
2480
Single Engines
Walter Titan/A1 – 25 4000
Table 2 Comparison of engines at Maximum Continuous Rating, 30000ft, ISA, 0.8 Mach
Net Thrust
lb
SFC
lb/hr/lb
Fuel Flow
lb/hr
GE 85 – 13
2160
1.285
2775
JT 15 D – 3
1741
0.846
1472
Larzac – 04
2000
0.973
1946
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
1770
0.942
1667
n/a
n/a
n/a
Twin Engines
Single Engines
Adour RT 172 – 06
Walter Titan/A1 – 25
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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
1.4
1/10 th scale first complete model for the Low Speed wind tunnel at Woodford
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
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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
Aspect ratio
Leading edge sweep
t/c ratio centreline
Fuselage length
Tailplane gross area
Tailplane aspect ratio
Leading edge sweep
Fin net area
Basic mass (Weapon
Trainer, 2 wing pylons)
5.287
23.8 deg
11.04 %
2
2
Taper ratio
18.51 m (199.24 ft )
0.34
t/c ratio tip
9.07 %
12,099 mm (39 ft 8 in)
2
2
Tail arm
3.72 m (40 ft )
3.55
34 deg
4,532 mm (14 ft 10 in)
Taper ratio
t/c ratio
2
2
3.135 m (33.75 ft ) Fin arm
3,477 kg (7,665 lb)
0.413
8.5 %
3,688 mm (12 ft 1 in)
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:
o
o
o
o
o
Front seat
Centreline, 15 downward;
Rear seat
Centreline, 7 downward; 40 port or starboard, 30 downward
o
40 port or starboard, 35 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 goahead 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
5g
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
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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.
o
There was a small increase in wing sweep to 26 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.
Paper No. 2013/01
Performance of HS 1182 and 1182A against ASR 397 (iss.2)
Item
ASR 397
Not more than 4,000 ft
25 kn Crosswind
Take-off and Landing
Performance
Approach 120-140 kn
Threshold 100-110 kn
0.85M at 30,000ft
Speeds - Maximum
- Max Continuous SL Not less than 420 kn
- Optimum Cruise
Not less than M = 0.6
Climb with full fuel, time
from brakes off to 30,000ft
Not more than 7min
Ceiling
Not less than 40,000 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
1.5.3
HS 1182
MET
MET
MET
112 kn
0.83
440 kn
0.7
8 min
7.2 min
60 % fuel
45,000 ft
5g
2.9 g
1.7 g
HS 1182A
MET
MET
MET
MET
0.85
MET
MET
Remarks
Limited by
gust ride
6.8 min
46,000 ft
5g
3g
2g
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.
18
Table 4
Type
Category
Propulsion
Equipment
standard
Length
Wing Span
Wing Area
HS 1182 A
Trainer / Strike
Adour
Full, with strike
provision
11.64 m (38.2 ft)
9.77 m (32.04 ft)
Paper No. 2013/01
Optional variants from HS 1182
HS 1182 V
Weapon Trainer
Viper
Reduced Standard
HS 1182 AT
Weapon Trainer
Adour
Reduced Standard
10.12 m (33.2 ft)
9.0 m (29.51 ft)
10.12 m (33.2 ft)
9.0 m (29.51 ft)
HS 1182 AJ
Trainer / Export
Adour
Intermediate
Standard
10.33 m (33.9 ft)
9.4 m (30.83 ft)
15.3m2 (164.7 ft2)
15.3m2 (164.7 ft2)
16.72 m2 (180 ft2)
18.02 m2 (194 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.
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
BAMPR - explain
Table 5
Feature
Max level speed at 30,000 ft, ISA,
Mach no
Max continuous level speed, sea level,
kn
Opt. range cruise speed at 30,000 ft,
Mach no
Approach speed, 60 % fuel, knots
Threshold speed, 60 % fuel, knots
Time to 30,000 ft from brakes off,
mins
Service ceiling, 60 % fuel, ISA,
Sortie “E” duration hours
Sustained g,
o
SL, ISA +15 C, 60 % fuel, 350 kn
60% fuel, 20,000 ft, ISA, 0.7 M
60% fuel, 35,000 ft, ISA, 0.7 M
Take-off distance to 15 m height, still
o
ft
air, full fuel, S.L., ISA + 15 C
Landing distance from 15 m height,
o
still air, full fuel, S.L., ISA + 15 C,
wet runway
ft
Performance of optional variants.
ASR 397
HS 1182 A
HS 1182 V
0.85
0.85
0.82
0.85
0.85
420
440
434
500
500
0.6 +
Met
0.6
0.65
0.7
120-140
100-110
Met
Met
125
104
127
106
130
110
7.0
6.8
8.5
5.9
7.0
13,400 m
40,000 ft
1.0
14,000 m
46,000 ft
1.0
14,300 m
47,000 ft
1.0
14,600 m
48,000 ft
1.0
14,300 m
47,000 ft
1.0
4.5
3.0
2.0
5.0
3.0
2.0
4.4
3.0
1.9
5.5
3.8
2.1
5.2
3.1
2.0
4,000
3,150
3,500
2,500
2,800
4,000
Met
3,870
3,900
4,000
19
HS 1182 AT HS 1182 AJ
1.6
Paper No. 2013/01
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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Paper No. 2013/01
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
from these manoeuvres.
forward
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
2.1
DEVELOPMENT OF THE HAWK T.Mk.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)
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
Average time
90
57.8
41
Action
To change or replace 15 equipment items
(e.g. control column handle, trim actuator)
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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
th
it became 9/16 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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Paper No. 2013/01
th
(b)
1/30 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.
th
The model was modified to become 1/28.5 scale but was little used thereafter, being
superseded by a later high speed model.
th
(d)
1/10 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) TwoDimensional 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.
th
(f)
A new 1/6 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.
th
(g)
A new 1/6 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.
rd
(h)
The intake and internal duct of the aircraft was very carefully designed, and a 1/3
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.
th
(i)
Two small (1/18 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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Paper No. 2013/01
(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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Paper No. 2013/01
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
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
26
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Paper No. 2013/01
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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Paper No. 2013/01
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 glassreinforced 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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Paper No. 2013/01
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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Paper No. 2013/01
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
with Matra rocket projectile packs
2.4
2.4.1
Figure 19
Foreground – Red Arrows
Hawk with centreline smoke fuel tank
Background – up to 88 Hawks, including
Red Arrows, could carry Sidewinders
Structural Testing
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.
34
Figure 21
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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
instrument layout. (Figure 22)
Figure 22
36
Hawk Front Cockpit Instruments
(key on following page)
Paper No. 2013/01
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.
37
Figure 23
Paper No. 2013/01
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.
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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.
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2.5.8
Paper No. 2013/01
Leading dimensions of the Hawk T.Mk.1.
Wing
Span
9.39 m
Reference area
2
16.69 m (179.6 ft )
Sweep at leading edge
26
Section thickness/chord ratio
At root
Dihedral angle
(30ft 10in)
2
o
10.9 %
2
o
Pylon centre lines
inboard
2.399 m (7ft 10in)
Fuselage overall length
11.459 m (37ft 7in)
Tailplane
Span
Reference area
4.39 m (14ft 5in)
5.284
Taper Ratio
0.34
Sweep at ¼ chord
23½
o
34.6
Section thickness/chord
ratio
8.5 %
9%
Wing setting
1
outboard
Tail Arm
Aspect Ratio
Dihedral angle
1.772 m (5ft 10in)
Reference area
2.508 m (27 ft )
2
45
2
o
o
3.422 m (11ft 3in)
4.3 m (14ft 1in)
4.45
0.33
Sweep at ¼ chord
Fin and Rudder
Height above fuselage
o
At extended tip
2
2
4.328 m ( 46.6 ft ) Taper Ratio
Aspect Ratio
30
o
o
- 10 (anhedral)
Fin arm
Section thickness/
chord ratio
3.492 m (11ft 5in)
Sweep at ¼ chord
39
8 % to 9 %
o
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)
2 crew
Fuselage fuel 841 litres (185 I.G.) at s.g = 0.79
Wing fuel
864 litres (190 I.G.) at s.g.= 0.79
Role equipment
3,450 kg
172 kg
664 kg
682 kg
249 kg
(7,606 lb)
(380 lb)
(1,462 lb)
(1,504 lb)
(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.
40
Rolling pull-outs
720 deg. roll
360 deg. roll
180 deg. roll
Paper No. 2013/01
1 “g”
0 to 3 “g”
-1 to 2.5 “g”
-2 to 5.3 “g”
200 to 500 kn or 0.81M
200 to 500 kn or 0.81M
200 to 425 kn
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.
Level speed at 2000 ft.
0.85 (0.81 Acce. Std.)
420 kn at S.L.(475 kn Acce.Std.)
0.87
521 kn.
Thrust boundary at 2000 ft. 350 kn.
at 20000 ft M=0.7
at 35000 ft. M=0.7
4 “g”
3 “g”
2 “g”
6 “g”
3 “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.
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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:
2
at M = 0.7 up to CL = 0.57
2
at M = 0.8 up to CL = 0.28
2
at M = 0.8 above CL = 0.28
CD = 0.0204 + 0.1079 CL
CD = 0.0204 + 0.1089 CL
or
CD = 0.0169 + 0.1528 CL
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
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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.
3.1
HAWK VARIANTS
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 fivepylon 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
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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
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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
X
X
X
X
X
X
Aden 30 mm Gun Pod
BR 125 125 kg bomb (free fall)
BR 250 250 kg bomb (free fall)
BRP 250 250 kg bomb (retarded)
BR 500 500 kg bomb (free fall)*
Mk.81 250 lb bomb (free fall)
Mk.81 SE 250 lb bomb (retarded)
Mk.82 500 lb bomb (free fall)
Mk.82 SE 500 lb bomb (retarded)
Mk.83 1000 lb bomb (free fall)*
540 lb MC bomb (free fall)
BL 755 Cluster bomb
Matra F155 M/N Rocket launcher
Matra F2 Rocket launcher
LAU 51 Rocket launcher
X
-X
X
-X
-X
X
X
----
X & XX
X
X
X
X
X & XX
X
X
X
X
X
X
X
X & XX
X
X
X
X
X & XX
X
X
X
X
X
X
X
Oerlikon Snora Rocket launcher
CBLS 100 Practice bomb carrier**
CBLS 200 Practice bomb carrier**
----
X
X
X
X
X
X
455 litre (100 IG) External fuel tank
----
X
X
X
----
Sidewinder AIM-9G
Matra Magic Air-air missile
---
X
X
---
Reconnaissance camera pod
Sea Eagle Air-surface missile
X
X
---
---
All are 14 inch twin suspension stores.
* Modified trailing edges to fins.
X indicates single carriage
XX indicates carriage on twin store carriers
** To carry free fall and retarded practice bombs
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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!
Figure 26
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.
Subsequent orders (1980)
were from Kenya for 12 aircraft,
and 20 for Indonesia.
Hawk Mk.60 series with heavy stores load
Eight 250 kg bombs and gun pod
Figure 27
Hawk ZA 101 with two AMRAAM,
two 130 imperial gallon tanks and Sea Eagle
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3.3
Paper No. 2013/01
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
47
Paper No. 2013/01
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.)
48
3.5
Paper No. 2013/01
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.
49
Paper No. 2013/01
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.
50
Paper No. 2013/01
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 resiting 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 twoseaters 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.
51
Paper No. 2013/01
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
52
Paper No. 2013/01
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.
53
4.328
46.6
2.508
27
(m2)
(ft2)
(m2)
(ft2)
Tailplane area
Fin area
Maximum weapon load
(kg)
(lb)
1,500
3,300
3,622
7,986
11.84
38.9
(m)
(ft)
Overall length
(kg)
(lb)
16.69
179.6
(m2)
(ft2)
Wing area
Weights
Empty
9.39
30.8
(m)
(ft)
Dimensions
Wing span
5,200
23,130
Nominal SLST, ISA
(lb)
(N)
Mk.151
T.Mk.1
Table 7
Adour Engine Type
3,000
6,613
2.508
27
4.328
46.6
11.84
38.9
16.69
179.6
9.39
30.8
54
5,340
23,750
Mk.851
Mk.50Series
3,000
6,613
4,012
8,845
2.508
27
4.328
46.6
12.43(Mk.67)
40.8
16.69
179.6
9.39
30.8
5,700
25,350
Mk..861
Mk.60 Series
Hawk Developments (up to 1995)
Paper No. 2013/01
3,000
6,613
4,400
9,700
2.61
28.1
4.328 (ex TCV)
46.6
12.43
40.8
16.69
179.6
9.94
32.6
5,845
26,000
Mk.871
Mk.100 Series
3,000
6,613
4,450
9,810
2.61
28.1
4.328 (ex TCV)
46.6
11.35
37.25
16.69
179.6
9.94
32.6
5,845 (6,030)
26,000 (27,000)
Mk.871
Mk.200 Series
(m)
(ft)
(m)
(ft)
(m)
(ft)
Service Ceiling
TO Ground Run (Clean)
Landing Ground Run
(10% fuel)
Max. level Mach no.
(knots)
(km/hour)
(kg)
(lb)
Maximum takeoff weight
Performance
Max. level speed
(kg)
(lb)
(kg)
(lb)
Maximum fuel (int.)
(s.g. 0.79)
(ext.)
Weights (continued)
-----
-----
14,600
48,000
535
990
0.87
5,700
12,570
1,346
2,963
---
T.Mk.1
-----
-----
15,200
50,000
535
990
0.88
7,350
16,200
55
1,346
2,963
932
2,055
Mk.50Series
550
1,800
710
2,330
14,000
46,000
545
1,009
0.88
9,100
20,060
1,304
2,876
932
2,055
Mk.60 Series
Table 7 (continued) Hawk Developments (up to 1995)
Paper No. 2013/01
605
1,980
640
2,100
13,550
44,500
540
1,000
0.88
9,100
20,060
750
o
2,470 (ISA +15 )
(with ‘chute)
640
2,100
13,700
45,000
540
1,000
0.88
9,100
20,060
1,304
2,876
932 (1,360 ferry)
2,055 (3,000 ferry)
Mk.200 Series
1,304
2,876
932 (1,360 ferry)
2,055 (3,000 ferry)
Mk.100 Series
---
60% fuel, 2,720 kg, 6,000 lb stores
T.Mk.1
+8 -4 (1,500 lb)
1,575
1,600+
With 2 x 130 IG ext. tanks
With 2 x 190 IG ext. tanks (flown but not cleared)
56
1,313
Clean
!,400+
1,360
125
As above but with 4 x 1000 lb bombs
Ferry range (nm)
345
Combat radius (nm) with gun pod, 2xAIM9, 4 x 500lb bombs, 2 x 130 imperial gallon external tanks
+3.9 g
+2g
(2 S/W) + 7.5 g
+4g
+6 -3
+8 -4
Mk.100 Series
60% int. fuel + 4 Sidewinders, Sea Level
20,000 ft
+6 -3
+8 -4
Mk.60 Series
+ 5.8 g
+3g
+6 -3
+8 -4
Mk.50Series
Maximum sustained turns
60% int. fuel, clean Sea Level
20,000ft
Maximum instantaneous turns
60% int. fuel + 2 or 4 Sidewinders Sea Level
20,000 ft
Paper No. 2013/01
Table 7 (continued) Hawk Developments (up to 1995)
Normal symmetric ‘g’
60% fuel, 1,360 kg, 3,000 lb stores
1,400+
1,365
+ 3.9 g
+2g
+ 5.9 g
+ 3.8 g
(4 S/W) + 7 g
+ 3.8 g
+6 -3
+8 -4
Mk.200 Series
3.8
Paper No. 2013/01
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
Country
Type
Abu Dhabi
Mk. 63
Hawk Orders and Deliveries (as of 2012)
Number
Mk. 63C
Mk. 102
Australia
Mk. 127
Bahrain
Canada
Mk. 129
Mk. 115
Dubai
Mk. 61,
Mk. 61A
Mk. 51
Mk. 51A
(Mk. 66
Finland
Dates Contract (C) or delivery
16
October 1984 to May 1985
(14 a/c converted to 63A and 2 of these to 63B)
4
February 1995 to March 1995
18
April 1993
34
April 2000 to October 2001
(including 1 test airframe)
6
October 2006 to December 2006
23
July 2000 to August 2004
(1 to Oman)
8
March 1983 to September 1983
1
June 1988
50
December 1980 to October 1985
7
November 1993 to September 1994
16
Purchased from Switzerland, not
included in total)
Table 8 (continued) Hawk Orders and Deliveries (as of 2012)
India
Indonesia
Kenya
Kuwait
Malaysia
Oman
Saudi Arabia
South Korea
South Africa
Switzerland
Mk. 132
Mk. 132
Mk. 132
Mk. 132
Mk. 53
Mk. 109
Mk. 209
Mk. 209
Mk. 52
Mk. 64
Mk. 108
Mk. 208
Mk. 103
Mk. 203
24
42
40
17
20
8
16
16
12
12
10
18
4
12
Mk. 65
30
Mk. 65A
20
Mk.165
22
(RAF T.Mk.2 standard)
Mk. 67
20
Mk. 120/LIFT 24
Mk. 66
20
November 2007 to November 2008
August 2008 onward
2010 (C)
Indian Navy
September 1980 to March 1984
May 1996 to March 1997
February1996 to March 1997
April 1999
April 1980 to May 1982
November 1985 to September 1986
January 1994 to September 1995
August 1994 to May 1995
December 1993 to January 1994
December 1994 to May 1995
August 1987 to October 1988
March 1997 to December 1997
Contract 23 May 2012
September 1992 to August 1993
May 2006 to 2008
November 1989 to November 1991
(WDS 2002)
57
Paper No. 2013/01
United Kingdom
T.Mk. 1
176**
28
United States*
Mk. 128
(T.Mk.2)
T-45A/C
TOTAL
994
Zimbabwe
Demonstrators, etc.
G-HAWK, ZA 101
ZJ 100
ZJ 951
ZG 200, ZH 200,
ZJ 201 RDA
November 1976 to February 1982
(** Contract document 175)
April 2009 onward
223
(incl.2 prototypes)
Mk. 60
8
Mk. 60A
5
Mk. 50
Mk. 100D
Mk.100 NDA
Mk. 200
Mk. 200
TOTAL
1
1
1
2
1
April 1988 to October 2009(* with McDonnell Douglas, now Boeing)
July 1982 to October 1982
June 1992 to September 1992
Rebuilt with Mk.102 extended nose.
New Development Aircraft
Radar Development Aircraft
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.
58
4.
4.1
Paper No. 2013/01
The Rolls-Royce Turbomeca (RRTM) Adour turbofan engine.
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.
59
4.2
Paper No. 2013/01
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.
60
ADOUR Mk.
Paper No. 2013/01
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
------
SL, ISA + 35˚C
4,160
4,400
5,120
5,850 (lbf)
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
99.7/100.7
102.5/101.5
108/100.2
(lbf)
At SL, ISA conditions.
N1/N2
(100 % N1 = 13,600 rpm,
SFC
0.71
0.74
101.7/101.9 (%)
100 % N2 = 15,512 rpm.)
0.78
0.76
(lb/hr/lb.th)
At M = 0.8, ISA, Sea Level, Maximum Rating.
Thrust
SFC
4,150
4,760
5,760
1.06
1.07
1.08
6,680 (lbf)
------ (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.
61
5
Paper No. 2013/01
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.
62
Figure 42
Figure 43
Paper No. 2013/01
Hawk ZA 101 carrying multiple rocket launchers overflies
HSA (Kingston) Airfield and Works at Dunsfold
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.”
63
January 1971
HSA publication HSK 51D
Paper No. 2013/01
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
Performance & Handling Report for C.A. Release
March 1976
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.
NATOPS Flight Manual Navy Model T-45A
20th October 2011
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.
64
Paper No. 2013/01
Low Speed
(a)
Time Scale
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
1/6th scale definitive complete model for Woodford
1973 / 75
(f)
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)
(b)
1/30th scale preliminary model and modifications, tested in
High Speed “Blow-Down” tunnel at Brough
Model (a) tested in the RAE 3 ft high speed wind tunnel
1971 / 72
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
(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)
(g)
(g)
1/6th scale rear fuselage modification for airbrake tests and
extension to M = 1.2 . ARA Bedford
1/6th scale RAF stores for (e) including store and pylon loads
1971 / 72
1973
1975 / 76
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.
65
APPENDIX 1.1 (b)
Paper No. 2013/01
HAWK T.Mk.1
FLIGHT TESTING HOURS.
Four development aircraft at BAe.
Two at A&AEE.
XX 154
XX 156
XX 157
XX 158
XX 159
XX 160
Handling, Flutter, Loads
Systems and Engine
Avionics and Weapons
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
700 approx
800 approx
SIGNIFICANT DATES
1st issue of HSA internal draft Spec.
November 1968
Official go-ahead
Specification agreed
October 1971
March 1972
First Flights
XX154
XX156
XX 157
XX 158
XX 159
XX 160
August 1974
May 1975
April 1975
July 1975
June 1975
November 1975
Initial C.A. Release (Flying Trainer)
2nd C.A. Release (Weapon Trainer)
1st delivery to RAF Valley
1st delivery to RAF Brawdy
1st Trainer RAF pilot solo
1st delivery, export aircraft (Finland)
October 1976
July 1977
November 1976
July 1977
July 1977
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.
66
APPENDIX 1.2
(a)
Paper No. 2013/01
COSTS
HAWK CASE STUDY
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.
67
Paper No. 2013/01
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
Raw material
£
609,000
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
APPENDIX 1.3
2.889,000
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
o
from the large flaps in their fully down (50 ) position. When the undercarriage was deployed,
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Paper No. 2013/01
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.
375 mm
outboard
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,
6
using the 9/16 scale half model of the Hawk. The Reynolds number for the tests was 2.4 x 10
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.
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Paper No. 2013/01
ȘT is tailplane angle
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 (-ȘT), the more stalled flow is seen, suggesting
o
the lift (downwards) is becoming limited. With ȘT still short of the full back stick position of -15 ,
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
o
o
off and tail on for ȘT at 0 and -12 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.
Looking at the broken line curves, for full flap vane and undercarriage up, it is clear that at about
o
o
cannotofbeattack
raisedofand
is a dive
increasing
steepness.
a positive
angle
ofo)attack
ȘT = not
- 12 o
so, pulling
theof
tail
from neutral
(0 ) to At
nearly
full back
(-12
does
angle
-10theorresult
the effectiveness
of the
tailplane returns
extent,
that there
is some pitching
nose-up moment
pitch to
increase
the tailplane
downwards
force at to
all,some
and there
is in
virtually
no positive
available
with up.
the stick
bring
the nose
This well
is theback.
consequence of the tail stall, and the ‘Phantom Dive’, where the nose
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
Ș = 0o
T
Angle of attack
70
Paper No. 2013/01
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
selectedcurves,
for the F4
Figure
A1.3.3
Pitching
moment
flaps fully
Phantom – but the extra drag of the open slot atdown
high speed
it wascurves)
not an option
onon
the(upper
Tailmeant
off (lower
and tail
o
Hawk.
curves) with tailplane angles of 0 and -12 o
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.
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Paper No. 2013/01
Tailplane vane at 0 o
o
Figure A1.3.4 Installation of Tailplane Canard Vane at 0 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 (- ȘT) at all angles of
attack in the usable range.
72
Paper No. 2013/01
ȘT = - 12
ȘT = - 8 o
ȘT = - 9
o
ȘT = - 8
o
o
ȘT = - 0
o
ȘT = - 11 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)
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
ȘT = - 14 o
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.
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Paper No. 2013/01
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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Paper No. 2013/01
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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Paper No. 2013/01
The technique adopted was to push the stick forward to achieve a nose down pitch angle of
o
about 10 degrees, then pull it sharply back to the back stop (ȘT of -15 ) 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
o
approach the manoeuvre at 130 knots IAS, tail at -5 and in level flight, pitch attitude zero.
o
The angle of attack is zero, with flap fully down, nominally at 50 . At 3½ seconds the stick is
o
pushed forward (ȘT = +5 ) and the pitch attitude starts to go nose down with the angle of attack
o
reaching approximately -13 . 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.
o
The aircraft responds immediately, the angle of attack reaching +9 even though the stick has
o
been returned to its earlier condition for ȘT = -5 . 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.
76
APPENDIX 1.4
Paper No. 2013/01
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
6
characteristics. The Reynolds number of the tests was fairly low, about 1.36 x 10 based on the
mean chord, at M=0.8, though the tests covered from M=0.45 to M=0.9. Flight Reynolds
6
number was about 11.5 x 10 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
o
(favourable) break at about 11 incidence. But when the tailplane was added, a distinct ‘rideo
up’ occurred at a lower incidence, about 10 , 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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Paper No. 2013/01
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
o
o
o
10 incidence, but the wing itself stalls at 11 , giving an overall pitch down at 12 . 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,
o
enter the low pressure region at a much higher incidence (7 to 8 ) 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 pitcho
up beyond about 9 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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Paper No. 2013/01
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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Paper No. 2013/01
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
o
at 10 , leading edges on the 25 % chord line, coupled with a rear row over the same region, but
o
centred on 65 % chord and starting 0.45 inches further outboard. These were toed out by 20 .
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
Station
“R”
“C”
“D”
“E”
Broken line – Vortex generators at 25% & 65% chord
Location
Just o/b of wing/fillet intersection
55% semispan
69% semispan
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
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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.
6.1
BACKGROUND AND HISTORY
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
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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
o
as 115 kt on a 90 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
o
had moderate sweep, about 24 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
o
about 60 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
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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
Aircraft procured
Fiscal year
Aircraft procured
1988
1989
1992
1993
1994
1995
1996
1997
1998
12
24
12
12
12
12
12
12
15
1999
2000
2001
2002
2003
2004
2005
2006
2007
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
March 1999
April 2000
First student class graduates.
200,000 fleet flying hours reached.
1,000th student graduated.
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April 2001
May 2001
September 2003
February 2008
October 2009
6.9
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First aircraft exceeds 5,000 hrs.
342,000 hours, 22,000 deck landings, 1,328 students graduated.
500,000 flight hours.
800,000 flight hours, 3,000 students graduated.
932,000 + hours, 3,600 students graduated.
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.
7.1
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LEADING PARTICULARS AND PERFORMANCE SUMMARY
Leading Particulars of the T-45A “Goshawk”
Wing span
Wing gross area
Wing sweepback on ¼ chord line
30ft. 10 in.
190 sq.ft.
23.7˚
(9.39m)
(7.66 sq.m.)
Stabilator (tailplane) span
Stabilator area
Stabilator sweepback, ¼ chord
15ft 1in
47.6 sq.ft.
30.1˚
(4.59m)
(4.43 sq.m.)
Vertical stabilizer (fin) area
Vertical stabilizer sweep, ¼ chord
Vertical stabilizer height
27.9 sq.ft.
39.5˚
6ft. 4in.
(2.59 sq.m.)
Operation ready weight, 2 crew, no fuel or pylons
(1.93 m.)
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)
Maximum take-off weight
13,340 lb (6050 kg)
up to 15,000 lb
Weight with 60% fuel (clean)
7.2
12,167 lb (5519 kg)
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
with take-off flap setting
with landing flap setting
Landing (estimated) in still air.
Approach Speed
3,420 ft
3,900 ft
at SL, ISA
at SL, ISA +15˚C
Approx. 80% of the above values.
11,000 lb
12,000 lb
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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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)
NASAR
Mean Altitude 39,200 ft.
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
(ft)
TAS Mach NASAR
(kn) number nautical miles
TAS Mach NASAR
(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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Paper No. 2013/01
Turning Performance.
12,500 lb, ISA, no pylons or stores
5,000 ft
15,000 ft
Max. sustained turn rate (deg/sec)
at Mach number
12.5
0.4
8.7
0.5
Max. sustained normal acceleration
at Mach number
Min. turn radius (buffet onset)
4.6 g
0.72
3.4 g
0.75
2,200 ft
3,200 ft
Lift /buffet limited turns:
7.5
Max. turn rate (deg/sec)
at Mach number
16.5
0.65
12.5
0.75
Max. normal acceleration
6.5 g
5.7 g
Maximum Level Speed.
Maximum thrust, ISA day.
13,000 lb gross weight, no pylons or stores.
Altitude (ft)
7.6
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
Structural and flight limits.
Max. normal acceleration, low level
10,000 ft and above
6.5 g
7.33 g
Max. negative normal acceleration
-4g
Max. speed
550 kn CAS
Max. Mach number
1.04
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APPENDIX 2.1
A2.1.1
Paper No. 2013/01
THE MINIMUM POWERED APPROACH SPEED VPA MIN
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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Handling
CL MAX
(Min Speed)
Aircraft
Flap
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
FSFV
Fence
587
3-5
1.845
623
Acceptable
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
3-4
2.125
b.s. added
638
Good
Acceptable
5
2.12
2 to 5
1.95 to 2.2
average 2.06
3-4
1.688
1.685
1.697
G-HAWK
Various arrays turbs
CBFV
RAF
Flight
No.
Buffet
margin
knots
Leading edge
devices
640 to Acceptable to
good
661
Good
667
692
Good
RAF
450
Acceptable
3-4
6-7
3 mini fences and
outer fence
455
Good
2-4
1.886
No outer fence
457
Good
3-4
1.90
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
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
XX 154
CBFV
CBFV
XX 154
CBFV
Notes:
XX 154 seemed to be particularly good. XX 338 was poor and needed rectification.
Flaps
CBFV
Cut back flap vane;
Leading edge devices
FSFV
Full span flap vane
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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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
A2.3.1
Paper No. 2013/01
TAILPLANE STALL ON THE T-45
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
Paper No. 2013/01
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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Paper No. 2013/01
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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Paper No. 2013/01
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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Paper No. 2013/01
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
Figure 49
Throttle Handle
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
Paper No. 2013/01
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 at L.E., 21.5 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)
o
o
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
2 crew
Full internal fuel
½ flap
Combat Role
1 crew
Full internal + external fuel
¾ flap
Weight
13,292 lb (6,027 kg)
17,793 lb (8,071 kg)
ISA (15˚C)
Ground Roll
Distance to 50 ft (15 m)
1,919 ft (585 m)
2,936 ft (895 m)
3,494 ft (1,065 m)
4,823 ft (1,470 m)
ISA + 20˚ C (35˚C)
Ground Roll
2,329 ft (710 m)
3,445 ft (1,050 m)
4,281 ft (1,305 m)
5,774 ft (1,760 m)
ISA + 35˚C (50˚C)
Ground Roll
2,838 ft (865 m)
4,052 ft (1,235 m)
5,298 ft (1,615 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
Full fuel
13,290 lb
(6,027 kg)
Weight
Combat Role, 1 crew
10% fuel
10,754 lb
(4,878 kg)
10% fuel, stores gone
12,209 lb
(5,538 kg)
ISA (15˚C)
Ground Roll
Distance from 50 ft (15 m)
2,871 ft (875 m) 2,100 ft (640 m)
6,512 ft (1,985 m) 3,625 ft (1,105 m)
2,493 ft (760 m)
4,101 ft (1,250 m)
ISA + 20˚ C (35˚C)
Ground Roll
3,068 ft (935 m) 2,231 ft (680 m)
4,544 ft (1,385 m) 3,625 ft (1,105 m)
2,641 ft (805 m)
4,101 ft (1,250 m)
ISA + 35˚C (50˚C)
Ground Roll
3,215 ft (980 m) 2,329 ft (710 m)
4,987 ft (1,520 m) 3,920 ft (1,195 m)
2,707 ft (825 m)
4,413 ft (1,345 m)
Optimum Fuel Flow
Sea level
10,000 ft (3,048 m)
20,000 ft (6,096 m)
30,000 ft (9,144 m)
Clean aircraft, 12125 lb 5500 kg, ISA
at M = 0.3, 200 kn TAS
at M = 0.35, 223 kn TAS
at M = 0.4, 246 kn TAS
at M = 0.5, 295 kn TAS
18.7 lb/min (8.5 kg/min)
Sustained Level Turns at Sea Level at optimum speed.
Weight
Clean Aircraft
12,160 lb (5,516 kg)
Combat Aircraft
13,618 lb (6,177 kg)
Without combat flap
ISA conditions
ISA + 20˚C
6.0 g
5.3 g
4.1 g
3.6 g
With combat flap
ISA
ISA + 20˚C
15.8 deg/sec
14.2 deg/sec
12.7 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.
Hawk type.
Mk.151
T.Mk.1
In-service date
1976
1988
2002
5,200
6,030
6,500
Sea Level Static Thrust (lbf)
By-Pass Ratio
Mk.871
Mk.60, 100,
& 200 series
0.79:1
0.76:1
Compressor pressure ratio
10.7
11.3
Air Mass Flow (lb/sec)
94
97.6
Mk.951
Mk.120 +
series
0.78:1
12.2
105
Specific Fuel Consumption
(lb/hr/lbf thrust)
0.71
0.78
0.78
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 LMKT 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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Paper No. 2013/01
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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Paper No. 2013/01
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 No. 2013/01
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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Paper No. 2013/01
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.
120

The HAWK Story - Royal Aeronautical Society

Transcription

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