Methods of Manufacturing Cadmium Telluride Thin Film Photovoltaic Devices

Abstract

A control system (50) for a turbopropeller engine (2) of an aircraft (1) having a gas turbine (11) and a propeller assembly (3) coupled to the gas turbine (11), the gas turbine (11) having a compressor (12) coupled to an air intake (13) and a temperature sensor (22) being arranged in the air intake (13) to measure the temperature of engine intake air and provide a sensed temperature (T1.sub.sens); the control system envisages: a compensation system (40) to receive the sensed temperature (T1.sub.sens) from the temperature sensor (22) and to add to the sensed temperature (T1.sub.sens) a compensation quantity (comp) to compensate for a delay introduced by the time constant (τ) of the temperature sensor (22) and generate a compensated temperature (T1.sub.comp); and a control unit (20) to perform engine control operations based on the compensated temperature (T1.sub.comp). In particular, the compensation quantity (comp) is calculated based on an ISA International Standard Atmosphere—temperature (T1.sub.ISA), which is determined as a function of an external pressure (P0) measured by a pressure sensor (35).

Claims

1-18. (canceled)

19. A control system for a turbopropeller engine of an aircraft having a gas turbine and a propeller assembly coupled to the gas turbine, the gas turbine having a compressor coupled to an air intake and a temperature sensor being arranged in the air intake to measure the temperature of engine intake air and provide a sensed temperature, the control system comprising: a compensation system configured to receive the sensed temperature from the temperature sensor and to add to the sensed temperature a compensation quantity to compensate for a delay introduced by the time constant of the temperature sensor and generate a compensated temperature; and a control unit configured to perform engine control operations based on the compensated temperature.

20. The system according to claim 19, wherein the compensation system comprises: a first calculation block configured to receive a signal indicative of aircraft altitude and to calculate an ISA—International Standard Atmosphere—temperature at the corresponding altitude; a second calculation block coupled to the first calculation block and configured to receive the ISA temperature and to generate the compensation quantity based on the ISA temperature; and an adder block having a first adding input coupled to the second calculation block to receive the compensation quantity, a second adding input coupled to the temperature sensor to receive the sensed temperature, and a sum output configured to provide the compensated temperature.

21. The system according to claim 20, wherein the first calculation block is coupled to a pressure sensor designed to measure an external pressure and is configured to receive the measured external pressure as the signal indicative of aircraft altitude informationand to calculate the ISA temperature as a function of the external pressure.

22. The system according to claim 20, wherein the second calculation block is configured to implement the following transfer function: ST/(1+ST), wherein t is the time constant of the temperature sensor.

23. The system according to claim 20, wherein the temperature sensor is configured to generate the sensed temperature starting from an actual temperature implementing a first order lag transfer function: 1/(1+ST), wherein t is the time constant of the temperature sensor.

24. The system according to claim 20, wherein the compensation system further comprises a saturation block configured to allow compensation only during the aircraft descent and to set the value of the compensation quantity to zero during the aircraft ascent.

25. The system according to claim 24, wherein the saturation block is configured to saturate the value of the compensation quantity to a maximum saturation value during the aircraft descent.

26. The system according claim 20, further comprising: an anti-noise filter block arranged between the first calculation block and the second calculation block configured to filter and limit the bandwidth of the signal indicative of aircraft altitude.

27. The system according to claim 20, wherein the second calculation block comprises: a derivative block implementing a derivative function of the ISA temperature according to the expression t({acute over (α)}T.Math.I3A/dt), wherein t is the time constant of the temperature sensor; and a first order lag block implementing a lag function according to the expression 1/(1+sx).

28. The system according to claim 19, wherein engine control operations performed by the control unit include controlling the position of a Variable Stator Vane—VSV—device coupled to the compressor to partialize air flowing to the compressor, so as to avoid a stall condition thereof, based on the compensated temperature.

29. A turbopropeller engine for an aircraft comprising: a gas turbine and a propeller assembly coupled to the gas turbine, the gas turbine having a compressor coupled to an air intake and a temperature sensor being arranged in the air intake to measure the temperature of engine intake air and to provide a sensed temperature; wherein the turbopropeller engine comprises the control system of claim 19.

30. An aircraft comprising the turbopropeller engine according to claim 29.

31. A control method for a turbopropeller engine having a gas turbine and a propeller assembly coupled to the gas turbine, the gas turbine having a compressor coupled to an air intake and a temperature sensor being arranged in the air intake to measure the temperature of engine intake air and to provide a sensed temperature, the control method comprising: receiving the sensed temperature from the temperature sensor and compensating the sensed temperature, adding to the sensed temperature a compensation quantity to compensate for a delay introduced by the time constant of the temperature sensor and generate a compensated temperature; and performing engine control operations based on the compensated temperature.

32. The method according to claim 31, wherein compensating comprises: receiving a signal indicative of aircraft altitude and calculating an ISA—International Standard Atmosphere—temperature (TliSA) at the corresponding altitude; generating the compensation quantity based on the ISA temperature; and providing the compensated temperature as the sum between the compensation quantity and the sensed temperature.

33. The method according to claim 32, wherein receiving comprises receiving a measure of an external pressure from a pressure sensor as the signal indicative of aircraft altitude information, and calculating comprises calculating the ISA temperature as a function of the external pressure.

34. The method according to claim 32, wherein generating the compensation quantity comprises implementing the following transfer function: ST/(1+ST), wherein t is the time constant of the temperature sensor.

35. The method according to claim 32, wherein generating the compensation quantity further comprises allowing the compensation only during the aircraft descent and setting the value of the compensation quantity to zero during the aircraft ascent.

36. The method according to claim 35, wherein generating the compensation quantity further comprises saturating the value of the compensation quantity to a maximum saturation value during the aircraft descent.

Description

[0023] For a better understanding of the present invention, preferred embodiments thereof are now described, purely as non-limiting examples, with reference to the attached drawings, wherein:

[0024] FIG. 1 is a perspective view of an aircraft provided with a turbopropeller engine;

[0025] FIG. 2 is a schematic block diagram of the turbopropeller engine of the aircraft;

[0026] FIG. 3 is a schematic diagram of a Variable Stator Vane (VSV) device of the turbopropeller engine;

[0027] FIG. 4 is a plot of quantities related to operation of the VSV device of FIG. 3;

[0028] FIG. 5 shows a delay model of an inlet temperature sensor in the turbopropeller engine;

[0029] FIGS. 6A-6B show plots highlighting the effects of a delay introduced by the inlet temperature sensor;

[0030] FIG. 7 is a schematic block diagram of a system for direct compensation of the lag introduced by the inlet temperature sensor;

[0031] FIG. 8 shows the plots of an actual temperature T1.sub.act and of a sensed temperature T1.sub.sens during the aircraft descent;

[0032] FIG. 9 is a Table according to International Standard Atmosphere (ISA) showing the variation of temperature and pressure at different altitude values;

[0033] FIG. 10 shows a plot of the ISA temperature T.sub.ISA as a function of the external pressure P0;

[0034] FIG. 11 shows a first embodiment of a compensation system according to the present solution;

[0035] FIG. 12 shows a second embodiment of a compensation system according to the present solution; and

[0036] FIG. 13 shows a further representation of the compensation system according to the present solution.

[0037] The present solution stems from the realization, by the Applicant that the temperature sensor 22 may introduce a substantial delay in the temperature sensing, due to the sensor's high time constant τ. For example, the high value of the time constant τ may be due to the sensor construction and/or to the sensor arrangement in the air intake 13 of the turbopropeller engine 2, which may be traversed by a limited flow of air during operation.

[0038] The delay introduced by the sensor time constant τ may be modelled by a first order system, as schematically shown in FIG. 5. In particular, the sensed temperature T1.sub.sens is schematically shown as being the result of a first order lag block 30, with transfer function 1/(1+sτ), applied to an actual temperature T1.sub.act, i.e. the real ambient temperature.

[0039] As shown in FIG. 6A, during the descent of the aircraft 1, the measure of the sensed temperature T1.sub.sens will thus be delayed by the sensor lag and will thus differ from the actual temperature T1.sub.act, causing a (negative) temperature difference between the sensed temperature T1.sub.sens and the actual temperature T1.sub.act.

[0040] As shown in FIG. 6B, this negative temperature difference will cause the VSV device 25 to track the nominal schedule with an open bias (being more open than required, i.e. being controlled with a lower beta angle β). When the VSV device 25 tracks the more open position, the compressor stability is reduced (the stall margin of the compressor 12 is lower) and an increased risk for compressor stall occurs.

[0041] As will be discussed in the following, an aspect of the present solution thus envisages a suitable compensation of the sensed temperature T1.sub.sens provided by the temperature sensor 22, in order to compensate for the delay introduced by the same temperature sensor 22.

[0042] As shown in FIG. 7, a possible compensation solution may envisage a direct compensation of the lag of the temperature sensor 22, by introducing a lead compensator block 32, cascaded to the first order lag block 30 (modelling the same temperature sensor 22, as discussed with reference to FIG. 5), wherein the lead compensator block 32 has a transfer function given by: (1+sτ)/(1+sτ.sub.a), τ.sub.a denoting a lead value.

[0043] By a proper choice of the lead value τ.sub.a, it may be possible to achieve a direct compensation of the lag on the sensed temperature T1.sub.sens, generating a compensated temperature T1.sub.comp.

[0044] The present Applicant, however, has realized that this solution suffers from a drawback due to the high value of the time constant τ, causing an amplification of the little variations introduced by the noise on the sensed temperature T1.sub.sens, reducing the useful signal.

[0045] In particular, if the difference between τ and τa is very high (since the value of the time constant τ is high), the lead compensator block 32 behavior is the same as a derivative block, thus amplifying the noise; this same noise deteriorates the quality of the signal, producing small variations on the VSV command.

[0046] Accordingly, an aspect of the present solution envisages the estimation of the delay and the error affecting the measure of the sensed temperature T1.sub.sens by the temperature sensor 22, in particular during the aircraft descent, using the measure of a different sensor (as will be discussed in the following, a sensor providing altitude information, in particular a pressure sensor) to produce a suitable compensation quantity.

[0047] In particular, assuming that the time constant τ of the temperature sensor 22 is known, it is possible to estimate the delay and the error affecting the temperature measure during the aircraft descent.

[0048] In detail, the dynamic error e(s) between the sensed temperature T1.sub.sens and the actual temperature Tl.sub.act can be expressed as:

e(s)=T1.sub.sens(s)−T1.sub.act(s)

[0049] Considering, as discussed above, the temperature sensor 22 as a first order lag model, the sensed temperature T1.sub.sens can be expressed as:

[00002]$T.Math.1_{s.Math.e.Math.n.Math.s}.Math.\left(s\right)=T.Math.1_{\mathrm{act}}.Math.\left(s\right).Math.\frac{1}{1+s.Math.\tau }$

so that the above expression for the error e(s) may be expressed as follows:

[00003]$e\left(s\right)=T.Math.1_{\mathrm{act}}.Math.\left(s\right).Math.\frac{-s.Math.\tau }{1+s.Math..Math.\tau }=-s.Math..Math.\tau .Math.T.Math..Math.1_{a.Math.c.Math.t}.Math.\left(s\right).Math.\frac{1}{1+s.Math.\tau }$

[0050] It is possible to estimate the error e(s) in steady state during a constant fast descent, i.e. considering the rate of descent as a constant value A (ramp descent).

[0051] Under this assumption, the actual temperature T1.sub.act can be expressed, in the time domain, as:

T1.sub.act(t)=At.Math.u(t)

and, in the Laplacian domain, as:

[00004]$T.Math.1_{a.Math.c.Math.t}.Math.\left(s\right)=\frac{A}{s^2}$

[0052] Considering the final value theorem:

[00005]$e\left(\infty \right)=\underset{t.fwdarw.\infty }{\lim }.Math..Math.e\left(t\right)=\underset{s.fwdarw.0}{\lim }.Math..Math.\mathrm{se}\left(s\right)=\underset{s.fwdarw.0}{\lim }.Math..Math.s.Math.\frac{A}{s^2}\left[\frac{-s.Math..Math.\tau }{1+s.Math..Math.\tau }\right]=-A.Math..Math.\tau$

[0053] Accordingly, the steady state error during a descent depends on the rate of descent A and the sensor lag (time constant τ).

[0054] Moreover, the rate of descent A can be expressed as:

[00006]$A=\frac{d.Math.T.Math.1_{\mathrm{act}}}{\mathrm{dt}}$

so that the above expression for the steady state error can also be formulated as:

[00007]$e\left(\infty \right)=-\frac{\mathrm{dT}.Math..Math.1_{\mathrm{act}}}{\mathrm{dt}}.Math.\tau .$

[0055] FIG. 8 shows the plots of the actual temperature T1.sub.act and of the sensed temperature T1.sub.sens during the aircraft descent time, highlighting the lag τ and the error introduced by the temperature sensor 22, according to what has been discussed above.

[0056] In order to determine the rate of descent A, the hypothesis can be made that the actual temperature T1.sub.act is given by the sum of the temperature calculated by ISA (International Standard Atmosphere) condition, i.e. the ISA temperature T1.sub.ISA, and an additional value ΔT.sub.day, associated to the day condition and that does not depend on the time of descent (in other words, it can reasonably be assumed that this additional value ΔT.sub.day is constant with time and as the altitude decreases):

T1.sub.act(t)=T1.sub.ISA(t)+ΔT.sub.day.

[0057] The ISA temperature T1.sub.ISA indeed increases as the altitude decreases, as shown by the detailed values reported in a Table according to the International Standard, shown in FIG. 9.

[0058] From the above expression (and considering that the additional value ΔT.sub.day is constant with time), it follows that:

[00008]$\frac{\mathrm{dT}.Math..Math.1_{a.Math.c.Math.t}}{d.Math.t}=\frac{\mathrm{dT}.Math..Math.1_{\mathrm{ISA}}}{\mathrm{dt}}=A;$$\mathrm{sT}.Math.1_{\mathrm{act}}=s.Math.T.Math.1_{I.Math.S.Math.A}=A.$

[0059] It follows that it is possible to calculate the descent rate A based on the knowledge of the ISA temperature T1.sub.ISA.

[0060] As shown in the same Table of FIG. 9, the ISA temperature T1.sub.ISA can be determined based on the altitude value, and the same altitude value can be determined as a function of the external pressure P0.

[0061] In this respect, FIG. 10 shows how the ISA temperature T.sub.ISA varies as a function of the external pressure P0, according to the values shown in the ISA table of FIG. 9.

[0062] The aircraft 1 is provided with a pressure sensor, denoted with 35 in the following (see FIG. 11), that is configured to sense the value of the external pressure P0, i.e. of the pressure in the environment external to the aircraft 1.

[0063] The error e(s), given, as discussed above, by the following expression:

[00009]$e\left(s\right)=-\mathrm{s\tau }.Math.T.Math..Math.1_{\mathrm{act}}.Math.\left(s\right).Math.\frac{1}{1+s.Math.\tau }$

can thus be compensated by adding, to the sensed temperature T.sub.sens, the following compensation quantity comp(s), equal and opposite to the above error expression:

[00010]$\mathrm{comp}\left(s\right)=\mathrm{s\tau }.Math.T.Math..Math.1_{\mathrm{ISA}}.Math.\left(s\right).Math.\frac{1}{1+s.Math.\tau }$

wherein the ISA temperature T.sub.ISA is determined as a function of the external pressure P0 measured by the pressure sensor 35:

T1.sub.ISA=f[P0]

(it is noted that the equivalence sT1.sub.act=sT1.sub.ISA has been exploited in the above expression).

[0064] In other words, adding the compensation quantity comp(s) to the sensed temperature T.sub.sens at the output of the temperature sensor 22 allows to fully compensate the error e(s) (i.e. e(s)=0).

[0065] FIG. 11 shows a compensation system, denoted with 40, according to a first embodiment of the discussed solution, exploiting the above discussed compensation solution.

[0066] The compensation system 40 includes:

[0067] a first calculation block 42, which receives at the input altitude information (in the example shown in FIG. 11, this altitude information is represented by the value of the external pressure P0 determined by pressure sensor 35), and calculates the ISA temperature T1.sub.ISA at the given altitude, in particular as a function of the external pressure P0, according to the above discussed expression T1.sub.ISA=f[P0] (based on the ISA table illustrated in FIG. 9);

[0068] a second calculation block 44, coupled to the first calculation block 42, which receives at the input the calculated ISA temperature T1.sub.ISA and generates the compensation quantity comp, implementing the transfer function sτ/(1+sτ) (corresponding to a derivative filter);

[0069] an adder block 45, having a first adding input coupled to the second calculation block 44 and receiving the compensation quantity comp, a second adding input receiving the sensed temperature T1.sub.sens (generated by the first order lag block 30 modelling the temperature sensor 22, with transfer function 1/(1+sτ), starting from the actual temperature T1.sub.act), and a sum output providing the compensated temperature T1.sub.comp (wherein the error e(s) has been compensated).

[0070] This compensated temperature T1.sub.comp, as shown schematically, can then be used by the electronic control unit 20 to implement control operations on the turbopropeller engine 2, such as for controlling the opening/closing state of the VSV device 25 as previously discussed in detail, thereby providing a complete control system 50.

[0071] FIG. 12 shows a further embodiment for the compensation system, again denoted with 40, which comprises, in addition to what is discussed above in connection with FIG. 11:

[0072] an anti-noise filter block 46, in particular a second order low pass filter, inserted between the first calculation block 42 and the second calculation block 44, configured to filter and limit the bandwidth of the signal indicative of the external pressure P0 provided by pressure sensor 35, so as to improve the signal-to-noise ratio and aliasing effect (also considering that, since the rate of descent is not very fast, the bandwidth required for the compensation signal is generally small); and

[0073] a saturation block 47, at the output of the second calculation block 44, configured to allow the compensation only during a descent of the aircraft 1.

[0074] During an ascent, the output of the saturation block 47 (representing the compensation quantity comp for the adder block 45) is always set to a 0 (zero) value, so no compensation is provided; during the descent, the output of the saturation block 47 is limited (saturated) to a saturation value T1.sub.sat, for example equal to 30K (this value can be tuned depending on the time constant τ), in order to determine a maximum value for the compensation quantity comp and avoid an over compensation.

[0075] FIG. 13 shows the same compensation system 40, wherein the second calculation block 44 is shown composed of:

[0076] a derivative block 44′, implementing a derivative function of the ISA temperature T.sub.ISA, according to the expression τ(dT1.sub.ISA/dt); and

[0077] a first order lag block 44″, implementing a lag function according to the expression 1/(1+sτ).

[0078] The advantages of the present solution are clear from the previous discussion.

[0079] In any case, it is again underlined that the present solution provides an effective system to compensate for the delay introduced by the temperature sensor 22, allowing to achieve improved engine control operations and reduced operating and maintenance costs.

[0080] Advantageously, the disclosed solution exploits altitude information provided by a different sensor to calculate a compensation quantity comp, that is used to compensate for the error on the sensed temperature T1.sub.sens measured by the temperature sensor 22 (thus implementing a “sensor fusion” algorithm). In particular, an external pressure sensor 35 is used to calculate the compensation quantity comp, the pressure sensor having a much quicker response (and a much lower time constant) than the temperature sensor 22.

[0081] Finally, it is clear that modifications and variations can be made to what is described and illustrated herein, without thereby departing from the scope of the present invention as defined in the appended claims.

[0082] In particular, it is underlined that, although generally applied to a fixed-wing aircraft, the present disclosure may further apply to rotary-wing aircraft, tilt-rotor aircraft, or other apparatuses including a pitch-changing propeller assembly and a gas generator coupled to the aircraft.