Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
1997-11-25
2001-02-27
Chin, Gary (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C340S963000, C073S17800T
Reexamination Certificate
active
06195598
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a power margin indicator for a rotary-wing aircraft, especially a helicopter.
2. Background Art
A helicopter is flown monitoring a great many instruments on the control panel, which instruments for the most part represent the operation of the engine and associated components and the aircraft. For physical reasons, there are many limitations that the pilot has to take into account at every moment during the flight. These various limitations generally depend on the type of flight and on the external conditions.
Most helicopters built these days are equipped with one or two turbine engines, generally with a free turbine. The power is therefore drawn off from a low-pressure stage of the turbine, which stage is mechanically independent of the compressor assembly and of the high-pressure stage of the turbine. As a turbine engine runs at between 30,000 and 50,000 revolutions per minute, a special reduction gearbox—the main gearbox (BTP)—is needed for connecting it to the rotor(s).
The thermal limitations on the engine and on the main gearbox allow three main regimes at which the engine is used to be defined:
the take-off regime that can be used for five to ten minutes and which corresponds to a level of torque on the gearbox and to heating up of the engine turbine which are permissible for a short period of time without causing appreciable damage; this is then known as the maximum take-off power (PMD),
the maximum continuous regime during which, at no time, are either the gearbox capabilities or the capabilities resulting from the maximum admissible heating past the high-pressure blading of the first turbine stage exceeded: this is the maximum continuous power (PMC),
the maximum transient regime, limited to one or two tenths of a second, sometimes protected by a governor stop: this is then known as the maximum transient power (PMT).
There are also emergency excess-power regimes in multi-engined aircraft, which are used if there is a breakdown of one engine:
the emergency regime during which the capabilities of the gearbox on the input stages and the thermal capabilities of the engine are used to their maximum: this is known as super-emergency power (PSU or OEI30″) which can be used for thirty consecutive seconds, at most, and three times during a flight. If OEI30″ is used, then the engine has to be taken out and overhauled;
the emergency regime during which the capabilities of the gearbox on the input stages and the capabilities of the engine are extensively used: this is then known as maximum emergency power (PMU or OEI2′) which can be used for two minutes after OEI30″ or for two minutes and thirty seconds consecutively, at most;
the emergency regime during which the capabilities of the gearbox on the input stages and the thermal capabilities of the engine are used without causing damage: this is known as intermediate emergency power (PIU or OEIcont) which can be used for from thirty minutes to two hours (depending on the engine) continuously for the remainder of the flight following the engine breakdown.
The engine test engineer through calculation or testing, establishes the curves of available power of a turbine engine as a function of the altitude and of the temperature, and does this for each of the six regimes defined hereinabove.
The limitations mentioned are generally monitored using three parameters: the speed of the gas generator (Ng), the engine torque (Cm) and the exhaust temperature of gases at the inlet to the free turbine (T4).
BROAD DESCRIPTION OF THE ENGINE
The purpose of the present invention is to provide a power margin indicator for a rotary-wing aircraft, especially a helicopter, that allows the pilot(s) to have a combined indication of the power margins, replacing a number of conventional indicators that are generally scattered across the control panel.
For this, the power margin indicator for a rotary-wing aircraft, especially a helicopter, intended to give information on the power margin available on at least one engine and one gearbox of the aircraft, as a function of the flight conditions, is noteworthy, according to the invention, in that it comprises:
input means for the various control parameters of the engine and of the gearbox,
calculation means connected to said input means which, on the basis of the values of the control parameters and the values of the limits for the various regimes in which the engine is used, formulate the power margin, expressed as a collective-pitch value, and
display means showing, on a display screen, said power margins represented on a scale which is graduated in collective-pitch equivalents, capable of moving past a collective-pitch index in a window of said display screen.
As a preference, the power margin is calculated in said calculation means with respect to the first of the limits that would be reached by one of the limiting parameters of the engine or of the gearbox should the power requirement vary, this being for each of the regimes in which the engine is used.
Advantageously, said power margin scale, which is arranged vertically, is graduated from 0 to 10 degrees of pitch.
In addition, said power margin scale may be bordered by colored bands that indicate the various limits on use depending on the type of flight.
Thus the power margin indicator according to the invention may show, at every minute and in a combined way, the following information:
the value of the pitch,
the pitch indications of the first limit for each of the regimes in which the engine is used.
Furthermore, it may be envisaged that the display differs depending on whether the flight conditions are normal or not and/or that the pitch limits displayed are filtered by anticipating the behavior of the engine and altering the calculation laws as a function of the phase of flight. Likewise, the introduction of tests for consistency between the control and calculated parameters used, and/or the displaying of alarms if the indicator's operation should become degraded or if the displayed limits are exceeded in terms of level and/or of duration, may be envisaged.
REFERENCES:
patent: 4034605 (1977-07-01), Green
patent: 4736331 (1988-04-01), Lappos et al.
patent: 5050081 (1991-09-01), Abbott et al.
patent: 5608627 (1997-03-01), Lecomte et al.
patent: 5886649 (1999-03-01), Francois
patent: 5908485 (1999-06-01), Germanetti
patent: 2 710 026 (1995-03-01), None
International (PCT) Published Patent Application No. WO96/26472.
Berlioz Raymond Jacques Gerard
Bosqui Olivier
Chin Gary
Eurocopter
Fisher Christen & Sabol
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