2014 IEEE International Conference on Power Electronics, Drives and Energy Systems (PEDES)
Fuzzy Logic Control of Modern Aircraft Electrical Power System during Transient and Steady-State Operating Conditions Reyad Abdel-Fadil, Ahmad Eid
Mazen Abdel-Salam
Electrical Engineering Department Aswan University Aswan, Egypt
[email protected] Abstract— In this paper, a modern civil Aircraft Electrical Power System (AEPS) is modeled, analyzed and controlled using Fuzzy Logic Control (FLC) technique. The FLC control provides a certain behavior for the AEPS during load switching to meet the military standards of the aircraft. FLC control technique used for Generator Control Unit (GCU), AC-DC Controlled rectifier, AC-AC inverter and DC-DC Converter. With this control, the voltage at the main buses (variable frequency bus – 270VDC bus – 28VDC bus-constant frequency bus) lies within standard limits at transient and steady state conditions. At the other hand, the generator frequency lies within standard limits for all the frequency operating range of 400-800Hz. The AEPS is tested using a simulation model in PSIM and the obtained results verify that the FLC is an efficient control technique for the variable frequency aircraft systems. Keywords— fuzzy logic control; modern aircraft; military standards; variable frequency; MEA
I.
INTRODUCTION
The conventional aircraft utilizes a combination of hydraulic, electrical, pneumatic and mechanical power transfer systems. Increasing use of electrical power is seen as the direction of technological opportunity for advanced aircraft power systems based on rapidly evolving technology advancements in power electronics, fault-tolerant electrical power distribution systems and electrical driven primary flight control actuator systems [1]. The concept of More Electrical Aircraft (MEA) implies increasing use of electrical power to drive aircraft subsystems that in the conventional aircraft, have been driven by a combination of mechanical, hydraulic and pneumatic systems. The objective of the MEA is to completely or partially replace the non-electrical power in the aircraft with electricity. This idea was first applied to meet the military for less overall weight of the aircraft, lower maintenance costs, higher reliability and better performance with increasing capacity and rating of the civil aircrafts [2]. The MEA concept is seen as the direction of aircraft power system technology in the future. The aircraft power system will employ multi-voltage level hybrid DC and AC systems. Thus, MEA electrical distribution systems are mainly in the form of multi-converter power systems. The electrical
978-1-4799-6373-7/14/$31.00 ©2014 IEEE
Electrical Engineering Department Assiut University Assiut, Egypt
[email protected] modern aircraft consists of four types of voltages; 400/200VAC(variable frequency), 28VDC and 270VDC [3]. The MEA is anticipated to achieve numerous advantages [4] such as optimizing the aircraft performance and decreasing the operation and maintenance costs. Moreover, MEA reduces the emissions of air pollutant gases from aircrafts, which can contribute in significantly solving some of the problems of climate change. However, the MEA puts some challenges on the aircraft electrical system, both in the amount of the required power, the processing and management of this power. In this paper, modeling, performance and control of an MEA electrical power distribution system is investigated. The MEA aircraft consists of multiple DC and AC buses at variable frequencies to accommodate with variable load types inside the aircraft. Fuzzy-logic-based control of all converters inside the aircraft system is applied. Fuzzy logic control (FLC) provides outstanding advantages over the conventional controllers (PID) such as easiness to develop, cover a wider range of operating conditions, and more readily customizable in natural language terms. The obtained results verify that the FLC is superior compared to conventional control for the modern aircraft electrical power system. Voltages at different buses as well as frequencies meet the aircraft standards [5] during both transient and steady state operating conditions. II.
MODERN AIRCRAFT ELECTRICAL POWER SYSTEM
The MEA electrical generation and conversion efficiencies are significantly higher than earlier non-MEA aircrafts. The improvement in efficiency is primarily due to the use of a variable frequency (VF) generator and advances in power electronics that allow much higher power conversion efficiencies. For instance, comparing the B787 to the B777, the efficiency measured at the AC output of the generator is 53% compared to 34% and at the ±270 VDC bus the efficiency is 51% compared to 25% [6], respectively. These low values of the efficiencies are due to the low efficiency of the mechanical engine. The Boeing 787 has an engine-mounted generator that produces a variable frequency 230 VAC output voltage. About 30% of the generated power is used directly (variablefrequency loads). The rest of the generated power is divided between the ±270 VDC bus loads using an autotransformer rectifier unit (ATRU) with an efficiency of 97%, the loads
connected at the 115 VAC 400 Hz equipped with a 98% efficient transformer and finally the loads connected to the 28 VDC bus equipped with a 80% efficient transformer rectifier unit (TRU). The studied electrical system of B787 [3, 6] specified at cruise condition is shown in Fig. 1.
C. B787 Electrical Loads Different kinds of loads exist inside the B787 distributed in the aircraft to provide the necessary needs and welfare of the passengers. The total load inside the B787 aircraft sums to around 1MW [7] as listed in table I. The total load connected to the 270 VDC bus represents about 43.2%, 3.4% are connected to the 28 VDC bus, whereas, about 27.2% and 18% are connected to the primary and secondary AC buses, respectively. The rest of 8.2% represents the total losses of the aircraft power system. This leads to an overall efficiency of 91.8% for the aircraft electrical system. In addition to the VF loads, constant frequency of 400 Hz loads also exist. In this case, an inverter is required to provide the constant frequency. TABLE I. Bus name
270 VDC Loads
Fig. 1. Boeing 787 electrical power system structure.
The structure and specifications of the studied B787 modern aircraft electrical power system consist of the following components. A. VF Synchronous Generators with their GCUs The modern B787 aircraft consists of four variable frequency three-phase synchronous generators, two at each side of the aircraft. Each of them has a total output power of 250 KVA with an output voltage of 400 VLL and the frequency ranges from 360 Hz to 800 Hz [7]. Another two auxiliary power units (APUs), with specifications similar to the main generators except for their output power of 225 kVA each, are exist. These APUs operate only in emergencies such as failure of one or more of the main generators. The studied generator information can be founded in reference [5]. Fuzzy logic control is used to provide a suitable excitation voltage for the GCU to keep the output voltage of the generator lies within the aircraft standards allowable limits [6] whenever the engine speed or electrical load changes. B. MEA Power Converter Units In B787 aircraft there are two AC buses (main or primary and secondary) to which related loads are connected. The aircraft generators are connected to the primary AC bus as well as the primary AC loads as shown in Fig. 1. A lower AC voltage of 115V/200V with a variable frequency of 380 to 800 Hz characterizes the secondary AC bus. In addition to the AC buses, there are another two DC buses with voltage ratings of 270 VDC and 28 VDC. To get the ratings of the secondary AC bus as well as the DC buses, different power converter units are used in the aircraft power system. These units include autotransformer unit (ATU), transformer rectifier unit (TRU) and autotransformer rectifier unit (ATRU) [7-8].
115 VAC Loads
28 VDC Loads
230 VAC Loads
III.
AIRCRAFT ELECTRICAL LOADS Type of loads
Rated Power (KW)
ECS/Pressurization
320
Hydraulics
40
Equip Cooling
40
ECS Fans
32
Ice Protection
60
Others
140
Flight Controls
14
Others
20
ICS Hydraulics
40
Galleys
120
Fuel Pumps
32
Forward Cargo AC
60
432
200
34
252
FUZZY LOGIC CONTROL APPROACH
Unlike conventional logic systems, the fuzzy logic control (FLC) is able to model inaccurate or imprecise models. The fuzzy logic approach offers a simpler, quicker and more reliable solution than conventional techniques [9-10]. The fuzzy logic controller has three main components. Fuzzification, which modifies crisp inputs (input values from real world) into linguistic variables to enable the input physical signal to use the rule-base through membership functions. Rule-base, where fuzzy inputs are compared and based on membership functions of each input. Defuzzification converts back the fuzzy outputs of the rule-base to crisp ones and selects membership functions for the different control outputs from the rule-base. FLC components shown in Fig. 2.
Fig. 2. Fuzzy logic control components.
A. Fuzzification Fuzzy logic uses linguistic variables instead of numerical variables. For example, in a control system, error between reference signal and output signal can be assigned as Negative Big (NB), Negative Medium (NM), Negative Small (NS), Zero (ZE), Positive Small (PS), Positive Medium (PM), Positive Big (PB). The triangular membership function is used for fuzzification. The process of fuzzification convert numerical variable (real number) to a linguistic variable (fuzzy number). NB
1
NM NS
ZE
PS
PM
PB
aircraft system, this structure is modified to get better performance and for some of them, PI-FLC is used. C. Defuzzification The rules of fuzzy logic generate demanded output in a linguistic variable, according to real world requirements, linguistic variables have to be transformed to crisp output (real number). The choices available for defuzzification are numerous. So far, the choice of strategy is a compromise between accuracy and computational intensity database. The membership functions used in this study for defuzzification are shows in Fig. 5. In defuzzification step, three different methods can be used such as center of area or centroid (COA), bisector, middle of maximum (MOM) [11]. The most popular method, the center of gravity or area is used for defuzzification as:
0.5 ∑ ∑
0 -4
-3
-2
-1
0
1
2
3
4 1
Fig. 3. Membership functions representing a first input (error).
B. Rule Evaluator Conventional controllers like PI and PID have control gains, which are numerical values. Fuzzy logic controller uses linguistic variables instead of the numerical values. The linguistic variables of error signal change of error signal, and output represents degree of membership functions. The basic fuzzy set operations needed for evaluation can use one of three rules of AND (∩), OR (∪) or NOT (~) [10]. In this work, AND-intersection is used as: (1)
μ A∩B = min [μ A (X), μ B (X)]
The rule base stores the linguistic control rules required by rule evaluator (decision-making logic). An example of the rules used in this work are listed in Table 2. It is worth mention here that the Table II is symmetrically diagonal about the membership function M (Medium) and every membership function is listed seven times.
E
CE CNL CNM CNS CZE CPS CPM CPL
NL VL VL VL VL LO UM M
TABLE II.
NM VL VL LO LO UM M AM
NS VL LO LO UM M AM AM
ZE
PS
LO LO UM M AM B B
UM UM M AM B B VB
(2)
PM UM M AM B B VB VB
VL
LO
UM
M
AM
B
VB
-2
-1
0
1
2
3
0.5 0 -4
-3
Fig. 4. Membership functions representing the output variable.
IV.
SIMULATION RESULTS AND DISCUSSIONS
The aircraft electrical power system shown in Fig. 1 composites from four identical channels, and hence, only one channel is simulated for simplicity and time saving in PSIM commercial software [12]. The FLC control is carried out using a C-code capability in PSIM software. Because of the lack of the engine model, a DC motor replaces it. The simulated single channel of B787 aircraft is shown in Fig. 6 showing its main buses and ratings. The FLC ensures that at every bus of the aircraft, the magnitude of the voltage and its frequency comply with the aircraft standards [6].
PL M AM B VB VB VB VB
IF-THEN RULE BASE FOR FUZZY LOGIC CONTROL
Moreover, this table is based on PD-FLC scheme. This structure works for the generator control unit, GCU, and, frequency control. For the power converters inside the studied
4
Fig. 5. One channel of the simulated B787 model.
The aircraft system is studied at severe conditions of a full load step at different frequency range of 400-800 Hz. In the following section, the voltage, frequency and power profiles are demonstrated at different buses. A. Generator Voltage and Frequency Control To meet the aircraft standards of the generator terminal voltage and frequency, the filed excitation voltage shaft speed and shaft speed are controlled as shown in Fig. 7. The shaft speed of the engine (DC motor equivalent) is controlled by changing the armature voltage keeping the excitation voltage constant. The terminal voltage of the DC motor is regulated by the PD-FLC to get the required speed and hence the frequency. By the same way, another PD-FLC controller is used to regulate the excitation voltage of the main synchronous generator yielding the required terminal voltage. The reference frequency 360-800 Hz, in this study three variable frequencies is used as case study (400-600-800) as show in fig.8.
power. The terminal voltage of the generator is shown in Fig. 12-a, and it is accepted compeered to standard limits. The waveforms of the voltage and currents are shown in Fig. 12-b. It is noted that the performance of voltage profile at 400 Hz and 600 Hz is better than at 800 Hz due to large variation of frequency and the design of FLC. The harmonic contents at 800 Hz are less than that at 400 Hz or 600 Hz. 300K 250K 200K 150K 100K 50K 0K
400.4 400.2 400 399.8 399.6 399.4 0.1
0.15
0.2 Time (s)
0.25
Fig. 7. Main generator output Power and frequency at 400Hz. 180
160
140
120
100
80 0.1
0.12
0.14
0.16
0.18
VOut −
+
Frref
+
Vref = 400V
0.2 Time (s)
0.22
0.24
0.26
0.28
0.3
(a)
−
Va
Vb
Vc
200 100 0 -100
Fig. 6. Shaft speed and filed voltage control with PD-FLC.
To test the control, the main generator is loaded with full load (250 KVA) at t = 0.1 sec and the system performance is analyzed at different frequency from 400 to 800 Hz at the following sections. 1) Operation at 400 Hz Applying the full load at 0.1 sec and at 400 Hz operation, the generator real/reactive power is shown in Fig. 7 with its frequency. The change in frequency is less than the standard value of ±1 Hz. The generator rms voltage value during transient and steady state conditions is shown in Fig. 8-a with the standard limits. It can be seen that the voltage lies satisfactorily within limits. The temporal waveforms of the generator three-phase voltages and currents are shown in Fig. 8-b. The existed harmonics in the waveforms are due to the switching actions of the power converters inside the system and there is no filter at the terminals of the generator. 2) Operation at 600 Hz With the same conditions at 400 Hz, the aircraft system is controlled to operate at 600 Hz. The generator frequency oscillates between 599.6-599.9 Hz, which is still within the limiting standards as shown in Fig. 9 with the generator power. The rms voltage value also lies inside the aircraft standard envelope as shown in Fig. 10-a and the temporal waveforms are shown in Fig. 10-b. 3) Operation at 800Hz The last case study is at a frequency of 800 Hz. The generator frequency oscillates between 799.4-799.5 Hz, which is acceptable variation as shown in Fig. 11 with the generator
-200 Ia
Ib
Ic
600 400 200 0 -200 -400 -600 0.2
0.201
0.202
0.203
0.204
0.205
Time (s)
(b) Fig. 8. Main AC Bus, 115 V, 300-800 Hz. (a)Voltage(rms) with aircraft stander (b) three phase voltage and current waveform. 300K 250K 200K 150K 100K 50K 0K
600 599.8 599.6 599.4 599.2 0.1
0.15
0.2 Time (s)
0.25
0.3
Fig. 9. Main generator output Power and frequency at 600 Hz.
180
160
MIL-STD
140
120
100
80 0.1
0.12
0.14
0.16
0.18
0.2 Time (s)
(a)
0.22
0.24
0.26
0.28
0.3
Va
Vb
Vc
S1 0.2
0.201
0.202
0.203
0.204
S6
Ic
S3
Ib
S5
Ia 600 400 200 0 -200 -400 -600
S4
0
-200
S2
100
-100
270 Dc Bus
3-Phase VF (360-800)HZ 230/400V
200
Vref = 270V +
0.205
Time (s)
(b) Fig. 10. Main AC Bus, 115V, 300-800 Hz. (a) Voltage(rms) with aircraft stander (b) three phase voltage and current waveform.
− VOut
Fig. 13.
∫
Three Phase PWM Controlled Rectifier Topology with PI-FLC
350
250K 325
200K 150K
MIL-STD
300
100K 50K
V_ out
275
0K 250
799.7
225
799.6 200
799.5 799.4
0.1
0.15
799.3
0.2 Time (s)
0.25
0.3
799.2 0.1
0.15
0.2 Time (s)
Fig. 14. 270VDCBus Voltage with stander limit.
0.25
Fig. 11. Main generator output Power and frequency at 800Hz.
C. 28 VDC Bus Voltage Control The 28VDC bus feeds the pilot cabin communication and electronic devices and it connected to batteries of emergency conditions. A three-phase bridge rectifier with forward buck converter is used to provide a 28VDC from 400VAC. The forward buck converter is controlled by PD-FLC type as shown in Fig. 15. The simulation results of the voltage at 28VDC bus at the same previous conditions are shown in Fig. 16.
180
MIL-STD
160
140
120
100
80 0.14
0.16
0.18 Time (s)
0.2
0.22
0.24
(a) Va
Vb
Vc
200 100 0 -100 -200 Ia
Ib
28 Vdc Bus
0.12
3-Phase VF (360-800)HZ 230/400V
0.1
Ic
Vref = 28v +
600 400 200 0 -200 -400 -600
VOut − D 0.2
0.201
0.202
0.203
0.204
0.205
Time (s)
(b) Fig. 12. Main AC Bus, 115V, 300-800 Hz. (a) Voltage(rms) with aircraft stander (b) three phase voltage and current waveform.
B. 270 VDC Bus Voltage Control The 270 VDC bus is considered a very important bus at MEA as about half of total load is connected at this bus. In this study a three phase PWM controlled rectifier, Fig. 13, is used instead of using conventional tap changing transformer. The inverter is controlled using PI-FLC type. The simulation results for 270DC bus voltage at transient and steady state when all loads are switched-on at same time are shown in Fig. 14.
Fig. 15.
Forward Buck Converter Topology With FLC.
50 45
MIL-STD
40 35
V_ out
30 25 20 15 0.1
Fig. 16.
0.15
0.2 Time (s)
0.25
28VDC Bus. (a) Bus Voltage with stander limit.
0.3
D. Constant Frequency-Constant Voltage Bus Control Major of MEA load is of DC and VF AC types such as primary AC loads anti-icing system and galley. In another hand, some of AC loads need constant frequency and constant voltage supply. In this case, a DC link PWM inverter is used to convert from VF (360-800 Hz) to CF of 400 Hz. The inverter is controlled by PI-FLC type. The PWM topology and control [13, 14] are shown in Fig. 17. The CF bus voltage (rms), three phase voltage and current waveforms are shown in Fig. 18.
Vref = 200
Fig. 17.
Parameter and bus VF AC CF AC 270VDC 28VDC
THD AND VOLTAGE RIPPLES OF THE STUDIED SYSTEM THDv % THDi % 400Hz 600Hz 800Hz 400Hz 600Hz 800Hz 12.9 13.8 14.14 10.3 9.5 8.66 10.8 ------5.9 -------------------
V.
∆V% ------1.2 0.7
CONCLUSIONS
In this paper, modern variable frequency aircraft electrical power system is modelled and controlled by using FLC technique. The FLC is used for the entire power converters inside the aircraft as well as the synchronous generator voltage magnitude and frequency. By adopting the FLC, the voltage magnitude at relevant buses comply with the aircraft standards during the transient and stead state periods. Both PD-FLC and PI-FLC are used to enhance the studied system performance compared to using classical PID control. The obtained simulation results show the effectiveness of the FLC method to keep the system performance at satisfied level at variable frequency (360-800 Hz) operating conditions.
+
VOut −
TABLE III.
∫
DC link PWM Topology With PI-FLC
REFERENCES
180
160
[1]
MIL-STD
140
V_ out
120
[2]
100
80 0.1
0.12
0.14
0.16
0.18
0.2 Time (s)
0.22
0.24
0.26
0.28
0.3
[3]
(a) va_CF
Vb_CF
Vc_CF
200
[4]
100 0
[5]
-100 -200 Ia_CF
Ib_CF
Ic_CF
200 100 0 -100 -200 0.2
0.202
0.204
0.206
0.208
0.21
[6] [7]
Time (s)
(b) Fig. 18. CF AC Load Bus.(a) Voltage(rms) with aircraft stander (b) Three phase voltage and current waveform.
The total harmonic distrotion (THD) and DC ripple voltage are calculated at all aircraft buses and listed in Table III for different frequency operations. It can be noticed that the THDv of the voltage at the VF bus increases with the frequency, while, the THDi of the phase current deceases with increasing the freqiuency due to the inductive load nature. At the CF bus, the THDv and THDi are calculated at a single frequency of 400Hz. All THD values are greater than the IEEE standard of 5%, which will be treated in future work. The DC voltage ripples for both 270V and 28V are less than the aircraft standard [6].
[8]
[9] [10] [11] [12] [13]
[14]
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