VSKYLABS Handley Page H.P.115
(This is a simple Plane-Maker model, with simple 2D cockpit panel).
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The H.P.115 was an aerodynamic research aircraft which made it's first flight in 1961.
It had a slender, low aspect ratio delta wing, and the engine was mounted above the rear of the fuselage at the base of the tail-fin. It's construction was all metal, except of the rudder and elevons, which were fabric covered.
For it's flight testing, the H.P.115 had a wing with leading edge sweep of 74 degrees. The leading edge was detachable to permit flight testing with a wide variety of shapes. A large airbrake was fitted under each wing, ahead of the main legs of the non-retractable landing gears. A camera was positioned on the fin, to photograph the tufts on the wing during flight testing. An anti-spin and braking parachute was located under the rudder.
The H.P.115 studied stability, control and handling characteristics which was aimed for the Concorde airliner development program.
Some more interesting facts:
- It was intended to be a glider, being towed to high altitude of around 30,000 ft.
- The fin had a bullet fairing at the top to accommodate a camera to record airflow experiments. Smoke generators mounted on the wing leading edges.
- The airfoil was a bi-convex type with the maximum thickness at 40% of the chord. This section was chosen as being representative of the type likely to be adopted for a supersonic transport.
- Crew: 1
- Length: 45 ft
- Wingspan: 20 ft (6.1 m)
- Height: 12 ft 9 in (3.9 m)
- Wing area: 430 ft²
- Airfoil: Bicon 6%
- Empty weight: 3,680 lb (1,670 kg)
- Loaded weight: 5,000 lb (2,291 kg)
- Powerplant: 1 × Bristol Siddeley Viper 9 turbojet, 1,900 lbf static.
- Maximum speed: 248 mph (399 km/h)
- Endurance: 40 minutes
H.P.115 Vortex Breakdown
The data below contained detailed information taken from materials which approved for unlimited public release.
Water tunnel studies I on models of highly swept wings with sharp leading edges have shown that at some position along the vortices associated with the flow past such wings a radical change in the nature of the flow can occur. The vortex expands radially and the line of the core takes on a spiral shape. The phenomenon is usually described as 'vortex breakdown', and the initial appearance is downstream of the wing. As incidence is
increased the position of breakdown moves upstream and may occur forward of the trailing edge at sufficiently large angles of incidence Because of the current interest in slender wings and in the turbulence in their wakes, it was decided to see if vortex breakdown could be identified in flight behind the HP 115 research aircraft. This Report describes the experimental technique used and the results obtained during tests in 1964 and 1965 at the Royal Aircraft Establishment, Bedford.
The HP 115 Aircraft
This aircraft was specifically designed and constructed to investigate the low speed handling problems of slender winged aircraft. It has a wing of triangular planform with rounded tips, a leading edge sweep of 76 degrees, and an aspect ratio of 0"92. The wing has a biconvex circular arc section with a constant thickness/chord ratio of 0"06, and the leading edges are effectively sharp, having a radius of 0.1 inch. Large full-span elevons are fitted, it being considered at the design stage that separate elevators and ailerons might introduce control problems if the vortex should cross the chordwise elevator-aileron boundary.
From the available wind tunnel data on models having a general resemblance to the HP115 wing planform (but not wing section shape) it was estimated that the incidence at which vortex breakdown would occur at the trailing edge would be about 35 degrees at an indicated airspeed in the region of 45-50 kn, although the performance of the aircraft was such that this speed would be associated with a high rate of descent. Angles of incidence of this order are well outside the capabilities of conventional aircraft, and at the time when the trials were planned, had not been reached by the HP 115. Previous flight experience with the aircraft at speeds down to about 60 kn had shown that it was remarkably docile with no insuperable handling problems, and as a preliminary to the flow visualization tests, a number of flights was made at progressively lower speeds: these showed that the aircraft remained fully controllable in the incidence and speed range required (35 degrees and 45 kn, indicated). It was found however, that there was a marked reduction in the turbulence threshold required to initiate the Dutch roll at these airspeeds and angles of incidence. It was possible to damp out this oscillation quite rapidly by forward movement of the control column, thus reducing incidence, but it meant that very calm conditions had to be chosen for the flow visualization trials.
Flow Visualization Technique:
In order to make the flow visible, colored smoke was injected into the air stream at the predicted position of the vortex core, close to the intersection of the wing leading edge with the fuselage side. The smoke generating system consisted of a chemical cartridge adapted from a marine distress signal, a tar trap, and a pipe to direct the smoke into the vortex core. The cartridge, which was ignited electrically using a switch in the cockpit, produced dense orange smoke for approximately 30 seconds. Longer duration cartridges (60 and 120 seconds) were tested but did not produce sufficiently dense smoke for photographic purposes.
A test firing of the cartridge before it was installed on the aircraft showed that the smoke was accompanied by a considerable quantity of soot and tarry material and it was thought advisable to remove as much as possible of these undesirable products of combustion to reduce contamination of the airframe and engine. The cartridge was therefore mounted so that the smoke from it first entered a can containing baffles which reversed the flow twice (Fig. 2). At the forward end of this can a one inch (internal) diameter pipe led the smoke
over the leading edge and into the vortext core. Some preliminary flight tests were required before a satisfactory location for the pipe exit was obtained. The trap removed an estimated 75 per cent of the tar, and as a first attempt, was considered reasonably satisfactory. The untrapped tar was, however, sufficient to cause some inconvenience, and for future experiments an improved design would be desirable. Possible modifications could be a larger number of baffles and an increase in length of the trap to promote cooling and condensation of the tar. The untrapped tar was deposited on the wing upper surface and also on the engine compressor blades. Removal of the deposit from the wing was facilitated by applying a thin coat of lanolin to the upper surface before each flight. The contamination of the engine was not entirely unexpected since it seemed inevitable that some denser particles of the smoke emission would escape from the vortex core and might find their way into the engine intake. The degree of contamination was however, greater than expected, and sufficiently serious to require a cleaning treatment after each flight.
Flight Test Technique:
On each flight one smoke canister was carried under each wing. For straight runs the cartridges were fired individually so that two airspeed conditions could be observed, while for the examination of the effects of side slip, both cans were fired together and side slip progressively increased and reduced. The resulting flow patterns were observed from a chase aircraft and photographed with a handheld 16 mm cine camera. For most of the flights the chase aircraft was an Auster AOP Mk 9 but on a few occasions a Whirlwind helicopter was employed. Most of the flights were made in the speed range 45-65 kn and in these conditions the rate of descent of the H.P. 115 was of the order of 1000 ft per minute (5 meters per second). Considerable skill on the part of the chase aircraft pilot was required to maintain a suitable observation position relative to the target aircraft.
In straight flight at high incidence no undue difficulty was experienced in flying the aircraft in spite of the lack of forward view. On several occasions pilots reported that turbulence had initiated the Dutch roll but that the degree of rolling could be limited by instinctive lateral control movements or by reducing incidence. The comment made after the flight in which the highest incidence was achieved (37 degrees) was "The aircraft seems to be stick fixed unstable at indicated air speeds below about 46 to 47 kn (approximately 36 degrees) and
it is correspondingly difficult to ensure both a stable airspeed and minimum stick input at the same time. Very still air is needed for runs at these speeds, as the slightest disturbance sets off the unstable Dutch roll. Despite these comments the aircraft remains easy to control, and instinctive corrections to the Dutch roll oscillations will limit the degree of rolling with no sensation of being near an aircraft limit of controllability."
Although on this flight the vortex breakdown was forward of the trailing edge, no effect was felt by the pilot. When side slip was applied, there was a marked deterioration in the handling; only small values of indicated side slip (approximately 5 degrees) could be achieved at incidences of about 30 degrees before encountering elevon buffet. On one flight the pilot commented as follows "On a dummy run (i.e. not filmed or recorded) at lower speed 48 kn--some evidence of flow breakdown was felt at the rear of the aircraft when slipping with maximum aileron. The aircraft in this condition was not steady and had a small pitching, rolling and yawing motion. The general feel of the aircraft was not pleasant." The incidence in this case would have been about 34 degrees, the side slip angle approximately 5 degrees, and the breakdown position was probably forward of the trailing edge.
- Vortex breakdown has been shown to occur in flight, and the general characteristics of the flow associated with such breakdown are similar to those observed on models in wind and water tunnels.
- The position of breakdown moves upstream with increasing incidence in straight flight and with increasing side slip at constant incidence.
- The relation between burst position and incidence derived from flight tests is consistent with that obtained in model tests; part of the difference between flight and model test results may be attributed to elevon deflection in the flight case.
- Occurrence of vortex breakdown within +0.1 root chord of the trailing edge in straight flight caused no increase in handling problems on the lip 115 aircraft. With vortex breakdown close to the trailing edge in side slip conditions, some deterioration in stability took place.
This information above was taken from the 'Vortex Breakdown - Some Observations in Flight on the HP 115 Aircraft' report, by L. J. Fennel / Aerodynamics Department, R.A.E., Farnborough, Hants.