scholarly journals Simulating actuator energy consumption for trajectory optimisation

2017 ◽  
Vol 232 (11) ◽  
pp. 2178-2192
Author(s):  
Michael Cooper ◽  
Craig Lawson ◽  
Amir Zare Shahneh
Keyword(s):  
Control System ◽  
Flight Control ◽  
High Speed ◽  
Electrical Energy ◽  
Guidance System ◽  
Control Surface ◽  
Aerodynamic Load ◽  
Actuation System ◽  
Heading Control ◽  
Trajectory Optimisation

This work aims to construct a high-speed simulation tool which is used to quantify the dynamic actuator power consumption of an aircraft in flight, for use within trajectory optimisation packages. The purpose is to evaluate the energy penalties of the flight control actuation system as an aircraft manoeuvre along any arbitrary trajectory. The advantage is that the approximations include major transient properties which previous steady state techniques could not capture. The output can be used to provide feedback to a trajectory optimisation process to help it compute the aircraft level optimality of any given flight path. The tool features a six degree of freedom dynamic model of an aircraft which is combined with low frequency functional electro-mechanical actuator models in order to estimate the major transient power demands. The actuator models interact with the aircraft using an aerodynamic load estimator which generates load forces on the actuators that vary as a function of flight condition and control surface demands. A total energy control system is applied for longitudinal control and a total heading control system is implemented to manage the lateral motion. The outer loop is closed using a simple waypoint following guidance system with turn anticipation and variable turn radius control. To test the model, a simple trajectory analysis is undertaken which quantifies a heading change executed with four different turn rates. The tool shows that the actuation system requires 12.8 times more electrical energy when performing a 90° turn with a radius of 400 m compared to 1000 m. A second test is performed to verify the model’s ability to track a longer trajectory under windy conditions.

Designing a Control Failure Survival System for High Speed Transport Aircraft Using Eigenvalue Assignment Method

2012 ◽  
Author(s):  
Zairil A. Zaludin
Keyword(s):  
Control System ◽  
Flight Control ◽  
High Speed ◽  
Control Surface ◽  
Longitudinal Motion ◽  
State Variables ◽  
Flight Control System ◽  
Short Period ◽  
Control Failure ◽  
Eigenvalue Assignment

Jika kerosakan berlaku kepada permukaan kawalan penerbangan, tujuan “Sistem Pereka Bentuk Kawalan Penerbangan” ialah untuk membahagi dan menyelaras usaha kawalan antara permukaan-permukaan kawalan yang masih aktif untuk tujuan mengekalkan mutu penerbangan yang diingini. Tugas utama ‘Sistem Pereka Bentuk Kawalan Penerbangan’ adalah untuk menyelaraskan unit kawalan semasa kerosakan berlaku ataupun menukar unit kawalan kepada sistem kawalan yang lebih sesuai untuk tujuan membaiki kerosakan tersebut. Dalam kertas ini, cara yang kedua dipertimbangkan. Reka bentuk “Sistem Keselamatan Kegagalan Kawalan” untuk pesawat hipersonik dibentangkan. Cara tersebut adalah berdasarkan cara penetapan nilai eigen dan teori pengatur kuadratik linear. Terdapat tiga masukan kawalan ke pesawat tersebut. Jika cara yang dibentangkan di dalam kertas ini digunakan, keputusan analisis yang dibentangkan menunjukkan bahawa sistem kawalan penerbangan untuk pesawat ini boleh direka bentuk sehingga kestabilan pesawat tersebut dicapai semula apabila salah satu ataupun gabungan permukaan kawalan gagal berfungsi pada masa yang sama. Didapati juga gerakan tabii pesawat yang mengalami kerosakan dapat dibaik pulih seperti sebelum kerosakan berlaku. Satu contoh disertakan dalam kertas ini menggunakan model matematik pesawat hipersonik. Kata kunci: dinamik penerbangan; penerbangan hipersonik; kawalan optimal; penetapan nilai eigen; Teori LQR In the event of a control surface failure, the purpose of a reconfigurable flight control system is to redistribute and coordinate the control effort among the aircraft’s remaining effective surfaces such that satisfactory flight performance is retained. A major task in control reconfiguration deals with adjusting the controller gains on-line or switching to a different control law to compensate for the failure. In this paper, the former option is considered. The design of a Control Failure Survival System (CFSS) for a hypersonic transport (HST) aircraft is presented. The method is based on eigenvalue assignment which was developed using Linear Quadratic Regulator theory. There are three control inputs available on board the HST; the change in the flaps deflection, the change in the propulsion diffuser area ratio and the change in the total temperature across combustor. Using the method discussed in this paper, the results showed that it was possible to reconfigure the flight control system such that the aircraft stability is regained when either a single or a combination of, control failures occurred simultaneously. In addition, the natural motion characteristics (i.e short period, phugoid and height motion) of the aircraft before the failure occurred are conserved and the transient response of the aircraft state variables after failure was almost the same as before failure occurred. An example is included in this paper using the mathematical model of the longitudinal motion of the HST. Key words: Aircraft dynamics; hypersonic flight; optimal control; eigen value assignment; LQR Theory

Design of sliding mode flight control system for a flexible aircraft

2021 ◽  
pp. 095441002110455
Author(s):  
Majeed Mohamed ◽  
Madhavan Gopakumar
Keyword(s):  
Control System ◽  
Rigid Body ◽  
Flight Control ◽  
High Speed ◽  
Sliding Mode ◽  
Flight Mechanics ◽  
Frequency Separation ◽  
Flight Control System ◽  
Flexible Aircraft ◽  
Aircraft Dynamics

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC.

scholarly journals Analisis Sistem Kendali Terbang dengan Gain Scheduling Pada Pesawat Efek Permukaan

WARTA ARDHIA ◽  
2017 ◽  
Vol 43 (2) ◽  
pp. 141
Author(s):  
Sayuti Syamsuar
Keyword(s):  
Control System ◽  
Flight Control ◽  
Surface Effect ◽  
Wind Tunnel Test ◽  
Longitudinal Mode ◽  
Gain Scheduling ◽  
Control Surface ◽  
Root Locus ◽  
Output Response ◽  
Adaptive Flight Control

This paper provides an overview of the design of adaptive flight control system of wing in surface effect craft Lippisch configuration 8 passengers capacity during cruise in the low speed and low altitude. The control system will be used the control surface, such as elevator deflection as input and pitch angle deflection as output response or by using engine throttle setting as input and others output response in the longitudinal mode. This paper describes some methodologies control system method and analysis such as PID controller system with gain scheduling approach, and root locus method. The observable matrices (4 x 4) on the longitudinal mode that used in the control system became from aerodynamic derivative parameters of 8 seaters configuration that calculated by DATCOM numerical simulation or wind tunnel test result and dummy data.  Kajian ini merupakan rancangan sistem kendali terbang adaptif pada pesawat efek permukaan konfigurasi Lippisch kapasitas 8 orang saat terbang mendatar pada kecepatan dan ketinggian terbang rendah. Sistem kendali terbang yang digunakan, seperti defleksi elevator sebagai input dan defleksi sudut pitch sebagai respon output atau penggunaan defleksi throttle mesin sebagai input dan parameter respon output lain pada gerak matra longitudinal. Kajian menjelaskan penggunaan beberapa metodologi dan analisis sistem kendali terbang adaptif, seperti kontroler PID dengan pendekatan gain scheduling, dan metoda root locus. Matriks ruang keadaan berukuran (4 x 4) pada matra longitudinal yang digunakan pada sistem kendali terbang adaptif diperoleh dari parameter turunan aerodinamika hasil perhitungan numerik DATCOM atau hasil uji terowongan angin dan data dummy.

RESPONSE OF A HIGH-SPEED WAVE-PIERCING CATAMARAN TO AN ACTIVE RIDE CONTROL SYSTEM

2021 ◽  
Vol 158 (A4) ◽  
Author(s):  
J AlaviMehr ◽  
M R Davis ◽  
J Lavroff ◽  
D S Holloway ◽  
G A Thomas
Keyword(s):  
Control System ◽  
Control Systems ◽  
High Speed ◽  
Control Surface ◽  
Control Algorithms ◽  
Active Centre ◽  
Ride Control ◽  
Control Step ◽  
Set Up ◽  
Ride Control System

Ride control systems on high-seed vessels are an important design features for improving passenger comfort and reducing motion sickness and dynamic structural loads. To investigate the performance of ride control systems a 2.5m catamaran model based on the 112m INCAT catamaran was tested with an active centre bow mounted T-Foil and two active stern mounted trim tabs. The model was set-up for towing tank tests in calm water to measure the motions response to ride control step inputs. Heave and pitch response were measured when the model was excited by deflections of the T-Foil and the stern tab separately. Appropriate combinations of the control surface deflections were then determined to produce pure heave and pure pitch response. This forms the basis for setting the gains of the ride control system to implement different control algorithms in terms of the heave and pitch motions in encountered waves. A two degree of freedom rigid body analysis was undertaken to theoretically evaluate the experimental results and showed close agreement with the tank test responses. This work gives an insight into the motions control response and forms the basis for future investigations of optimal control algorithms.

scholarly journals A new active asymmetry monitoring and control technique applied to critical aircraft flap control system failures

2019 ◽  
Vol 304 ◽  
pp. 04011
Author(s):  
Dario Belmonte ◽  
Matteo Davide Lorenzo Dalla Vedova ◽  
Gaetano Quattrocchi
Keyword(s):  
Control System ◽  
Flight Control ◽  
Angular Position ◽  
Control Technique ◽  
Monitoring And Control ◽  
Active Monitoring ◽  
Actuation System ◽  
Left And Right ◽  
And Control ◽  
Current Monitoring

Asymmetry limitation requirements between left and right wing flap surfaces play an important role in the design of the implementation of the secondary flight control system of modern airplanes. In fact, especially in the case of sudden breaking of one of the torsion bars of the flap transmission line, the huge asymmetries that can rapidly develop could compromise the lateral-directional controllability of the whole aircraft (up to cause catastrophic occurrences). Therefore, in order to guarantee the aircraft safety (especially during take-off and landing flight phase in which the effects of asymmetries could generate uncontrollable aircraft attitudes), it is mandatory to timely detect and neutralize these asymmetries. The current monitoring techniques generally evaluate the differential angular position between left and right surfaces and, in most the events, limit the Flaps Control System (FCS) asymmetries, but in severe fault conditions (e.g. under very high aerodynamic loads), unacceptable asymmetries could be generated, compromising the controllability of the aircraft. To this purpose, in this paper the authors propose a new active monitoring and control technique capable of detecting the increasing angular error between the different flap surfaces and that, after stopping the whole actuation system, acts on the portion of the actuation line still connected to the PDU to minimize the FCS asymmetries.

scholarly journals High Speed Microactuators for Low Aspect Ratio High Speed Micro Aircraft Surfaces

Actuators ◽  
2021 ◽  
Vol 10 (10) ◽  
pp. 265
Author(s):  
Ronald Barrett-Gonzalez ◽  
Nathan Wolf
Keyword(s):  
Aspect Ratio ◽  
Flight Control ◽  
High Speed ◽  
High Performance ◽  
Dynamic Range ◽  
Feedback System ◽  
Analytical Models ◽  
Control Surface ◽  
Operational Environment ◽  
Low Aspect Ratio

This paper covers a class of actuators for modern high speed, high performance subscale aircraft. The paper starts with an explanation of the challenges faced by micro aircraft, including low power, extremely tight volume constraints, and high actuator bandwidth requirements. A survey of suitable actuators and actuator materials demonstrates that several classes of piezoceramic actuators are ideally matched to the operational environment. While conventional, linear actuation of piezoelectric actuators can achieve some results, dramatic improvements via reverse-biased spring mechanisms can boost performance and actuator envelopes by nearly an order of magnitude. Among the highest performance, low weight configurations are post-buckled precompressed (PBP) actuator arrangements. Analytical models display large deflections at bandwidths compatible with micro aircraft flight control speed requirements. Bench testing of an example PBP micro actuator powered low aspect ratio flight control surface displays +/−11° deflections through 40 Hz, with no occupation of volume within the aircraft fuselage and good correlation between theory and experiment. A wind tunnel model of an example high speed micro aircraft was fabricated along with low aspect ratio PBP flight control surfaces, demonstrating stable deflection characteristics with increasing speed and actuator bandwidths so high that all major aeromechanical modes could be easily controlled. A new way to control such a PBP stabilator with a Limit Dynamic Driver is found to greatly expand the dynamic range of the stabilator, boosting the dynamic response of the stabilator by more than a factor of four with position feedback system engaged.

Reconfigurable Flight Control System Design Using Input-Output Information

2005 ◽  
Vol 219 (4) ◽  
pp. 277-285 ◽  
Author(s):  
S Hyung ◽  
Y Kim
Keyword(s):  
Adaptive Control ◽  
Control System ◽  
Flight Control ◽  
Fault Tolerant ◽  
Fault Tolerant Control ◽  
Control Surface ◽  
Control Technique ◽  
Input Output ◽  
Discrete State ◽  
Output Information

An adaptive control algorithm using input-output information is proposed for designing an aircraft fault tolerant control system. An input-output model is derived on the basis of a discrete state-space system. The formulated input-output model has the same structure as the autoregressive moving average (ARMA) model does, and therefore, the conventional system identification method using recursive least square can be used to identify the system. To design a reconfigurable control system, an LQ tracker with output feedback scheme is adopted. During the recursive adaptive control process, the system model is updated periodically. The proposed algorithm is applicable to time-varying systems in real time. To validate the performance of the proposed adaptive fault tolerant control technique, numerical simulation of the high performance aircraft with control surface damage was performed.

Development of a proof–of–concept aircraft smart control system

2009 ◽  
Vol 113 (1147) ◽  
pp. 587-590 ◽  
Author(s):  
P. Hutapea ◽  
K. Jacobs ◽  
M. Harper ◽  
E. Meyer ◽  
B. Roth
Keyword(s):  
Shape Memory Alloy ◽  
Shape Memory ◽  
Flight Control ◽  
Cooling System ◽  
Experimental Testing ◽  
Phase Changes ◽  
Control Surface ◽  
Control Mechanisms ◽  
Proof Of Concept ◽  
Actuation System

Abstract Hutapea et al showed that an actuation system based on shape memory alloy coils could be employed for a wing flap of an aircraft. A continued research and development of these previously demonstrated smart flight control mechanisms was performed with the goal to develop a proof-of-concept shape memory alloy (SMA) actuation system, which utilises SMA springs to control the six degrees of freedom of an aircraft. For this actuation system, the springs are heated via an electric current, causing the spring to contract as the metal’s phase changes from martensite to austenite. The contraction allows the springs to function as linear actuators for the aircraft’s control surfaces, specifically the flaps and ailerons on the wings and horizontal stabilisers and a rudder on the tail. As a significant advancement to the overall actuation system, an air burst-cooling system increases the cooling rate of the coils by means of forced convection. Computer-based finite element model analysis and experimental testing were used to define and optimise SMA spring specifications for each individual control surface design. A onesixth scale proof-of-concept model of a Piper PA-28 Cherokee 160 aircraft was constructed to demonstrate and to verify the final actuation system design.

Modeling, control design, and influence analysis of catapult-assisted take-off process for carrier-based aircrafts

2017 ◽  
Vol 232 (13) ◽  
pp. 2527-2540 ◽  
Author(s):  
Ziyang Zhen ◽  
Ju Jiang ◽  
Xinhua Wang ◽  
Kangwei Li
Keyword(s):  
Control System ◽  
Flight Control ◽  
Control Design ◽  
Control Surface ◽  
Flight Control System ◽  
Influence Analysis ◽  
Surface Ship ◽  
Lateral Channel ◽  
And Performance ◽  
Longitudinal Channel

This paper addresses the problems of modeling, control design, and influence analysis of the steam catapult-assisted take-off process of the carrier-based aircrafts. The mathematical models of the carrier-based aircraft, steam catapult, landing gears, and the environmental factors including deck motion and bow airflow have been established to express the aircraft dynamics in the take-off process. An engineering method based automatic flight control system has been designed, which is divided into the longitudinal channel and lateral channel. The influences of the preset control surface, ship deck motion, ship bow airflow, and automatic flight control system system are tested by a series of simulations. The simulation results show that the elevator angle preset is necessary in the stage of accelerated running on the ship deck and the deck motion is the most important factor for safe take-off, while the ship bow airflow is beneficial for climbing up of the aircraft. The automatic flight control system gives the guarantee of safety and performance in the take-off process of the carrier-based aircraft.

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