Trim Points Research Articles

Experimental Investigation of Propeller Induced Flow on Flying Wing Micro Aerial Vehicle for Improved 6DOF Modeling

In this research effect of propeller induced flow on aerodynamic characteristics of low aspect ratio flying wing micro aerial vehicle has been investigated experimentally in subsonic wind tunnel. Left turning tendencies of right-handed propellers have been discussed in literature, but not much work has been done to quantify them. In this research, we have quantified these tendencies as a change in aerodynamic coefficient with a change in advance ratio at a longitudinal trim angle of attack using subsonic wind tunnel. For experimental testing, three fixed pitch propeller diameters (5 inch, 6 inch and 7 inch), three propeller rotational speeds (7800, 10800 and 12300 RPMs) and three wind tunnel speeds (10, 15 and 20 m/s) have been considered to form up 27 advance ratios. Additionally, wind tunnel tests of 9 wind mill cases were conducted and considered as baseline. Experimental uncertainty assessment for measurement of forces and moments was carried out before conduct of wind tunnel tests. Large variation in lift, drag, yawing moment and rolling moment was captured at low advance ratios, which indicated their significance at high propeller rotational speeds and large propeller diameters. Side force and pitching moment did not reflect any significant change. L/D at trim point was found a nonlinear function of propeller diameter to wingspan ratio D/b, and propeller rotational speed. Rate and control derivatives were obtained using unsteady vortex lattice method with propeller induced flow effect modeled by Helical Vortex Modeling approach. In this research, we have proposed improved 6-DOF equations of motion, with a contribution of advance ratio J. It is concluded, that propeller induced flow effects have a significant contribution in flight dynamic modeling for vehicles with large propeller diameter to wingspan ratio, D/b of 22% or more.

Investigating Impaired Aircraft's Flight Envelope Variation Predictability Using Least-Squares Regression Analysis

Aircraft failures alter the aircraft dynamics and cause the maneuvering flight envelope to change. Such envelope variations are nonlinear and generally unpredictable by the pilot, as they are governed by the aircraft’s complex dynamics. Hence, to prevent in-flight loss of control, it is crucial to practically predict the impaired aircraft’s flight envelope variation due to any a priori unknown failure degree. This Paper investigates the predictability of the number of trim points within the maneuvering flight envelope and its centroid using both linear and nonlinear least-squares estimation methods. To do so, various polynomial models and nonlinear models based on the hyperbolic tangent function, which incorporate the influencing factors on the envelope variations as the inputs and estimate the centroid and the number of trim points of the maneuvering flight envelope at any intended failure degree, are developed and compared. Results indicate that both the polynomial and hyperbolic tangent function-based models are capable of predicting the impaired fight envelope variation with good precision. Furthermore, it is shown that the regression equation of the best polynomial fit enables direct assessment of the impaired aircraft’s flight envelope contraction and displacement sensitivity to the specific parameters characterizing aircraft failure and flight condition.

Design of a roll autopilot for a skid-to-turn guided missile

the design of autonomous systems is a growing research field which attract an increasing interest in various civilian and military applications. The design of control system can be considered a leading factor for developing any autonomous system. In this paper devoted to a missile autopilot which should achieve specific characteristics and performance criteria to ensure the overall system stability to track the desired commands in the presence of high dynamics and different uncertainties with high accuracy. The investigation of accurate missile 6DOF model is developed and linearization has been performed for designing the proper controllers for the pitch and roll channels and choose the optimum trim points based on system dynamic pressure. A design of lead compensator classical controller is presented to meet certain requirement of both time and frequency characteristic to ensure the system stability. The designed autopilot is evaluated using Matlab simulation is carried out for evaluating the proposed missile autopilot controller performance, the proposed controller presented sufficient performance during flight and simplicity, reliability, and availability for implementation in real time applications.

A method of partially overlapping point clouds registration based on differential evolution algorithm.

3D point cloud registration is a key technology in 3D point cloud processing, such as 3D reconstruction, object detection. Trimmed Iterative Closest Point algorithm is a prevalent method for registration of two partially overlapping clouds. However, it relies heavily on the initial value and is liable to be trapped in to local optimum. In this paper, we adapt the Differential Evolution algorithm to obtain global optimal solution. By design appropriate evolutionary operations, the algorithm can make the populations distributed more widely, and keep the individuals from concentrating to a local optimum. In the experiment, the proposed algorithm is compared with existing methods which are based on global optimization algorithm such as Genetic Algorithm and particle filters. And the results have demonstrated that the proposed algorithm is more robust and can converge to a good result in fewer generations.

Global Singularity-Free Aerodynamic Model for Algorithmic Flight Control of Tail Sitters

This paper addresses fundamental issues in tail-sitting and transition flight aerodynamics modeling in view of sum-of-squares (SOS) algorithmic guidance and control design. A novel approach, called theory, for modeling aerodynamic forces and moments is introduced herein. It yields polynomial-like differential equations of motion that are well suited to SOS solvers for real-time algorithmic guidance and control law synthesis. The proposed theory allows for first principles model parameter identification and captures dominant dynamical features over the entire flight envelope. Furthermore, theory yields numerically stable and consistent models for 360 deg angles of attack and sideslip. Additionally, an algorithm is provided for analytically computing all feasible longitudinal flight operating points. Finally, to establish -theory validity, predicted trim points and wind-tunnel experiments are compared.

Numerical study of saddle-node bifurcation for longitudinal flight with CFD/RBD technique

The square cross section missile was found instability in the longitudinal direction at a certain Mach number in a range of angle of attack both in CFD and wind tunnel test results. To deeply understand the flight behavior and the aerodynamics of the missile, the coupled Computational Fluid Dynamic/Rigid Body Dynamics (CFD/RBD) approach was employed to simulate the free flight process, and the predict technique of dynamic stability derivatives was founded to analyze the types of equilibrium points. The study shows that the type of trim angle of attack is turned from stable node to an unstable saddle with the increase of Mach number, known as the saddle-node bifurcation. There are two extra stable nodes are produced and located on the left and right sides of the saddle point when the Mach number is bigger than the transition Mach number. Further simulation results of free pitching motions reveal the decrease of the attraction region of the two extra stable modes, which makes the pitching motion more complicated and sensitive. A slight change of the initial angle of attack may make the missile move from one trim point to another.

Thermal Management of Single- and Dual-Tank Fuel-Flow Topologies Using an Optimal Control Strategy

The effect of fuel topology and control on thermal endurance of aircraft using fuel as a heat transfer agent was studied using an optimal dynamic solver (OPT). The dynamic optimal solutions of the differential equations governing the heat transfer of recirculated fuel flows for single- and dual-tank arrangements were obtained. The method can handle sudden jumps of operating conditions across different operating zones during mission and/or situations when control parameters have reached their physical limits. Although this method is robust in providing an optimal control strategy to prolong thermal endurance of aircrafts, it is not ideal for practical application because the method required iterative procedures to solve expensive nonlinear equations. The linear quadratic regulator (LQR), the feedback controller, can be derived by linearizing the adjoint equations at trim points to offer a simple control strategy, which can then be implemented directly in the feedback control hardware. The solutions obtained from both OPT and LQR were compared, and it was found two solutions were almost identical except in regions having sudden jump of operation conditions. Finally, a comparison between single- and dual-tank arrangements was made to demonstrate the importance of the flow topology. The study shows the dual-tank arrangement allows flexibility in how energy is managed and can release energy faster than a single-tank topology and hence provides improved aircraft thermal endurance.

Sum-of-Squares-Based Region of Attraction Analysis for Gain-Scheduled Three-Loop Autopilot

A conventional method of designing a missile autopilot is to linearize the original nonlinear dynamics at several trim points, then to determine linear controllers for each linearized model, and finally implement gain-scheduling technique. The validation of such a controller is often based on linear system analysis for the linear closed-loop system at the trim conditions. Although this type of gain-scheduled linear autopilot works well in practice, validation based solely on linear analysis may not be sufficient to fully characterize the closed-loop system especially when the aerodynamic coefficients exhibit substantial nonlinearity with respect to the flight condition. The purpose of this paper is to present a methodology for analyzing the stability of a gain-scheduled controller in a setting close to the original nonlinear setting. The method is based on sum-of-squares (SOS) optimization that can be used to characterize the region of attraction of a polynomial system by solving convex optimization problems. The applicability of the proposed SOS-based methodology is verified on a short-period autopilot of a skid-to-turn missile.

Helicopter flight simulation trim in the coordinated turn with the hybrid genetic algorithm

Helicopter trim models are multivariate nonlinear equations and it is difficult to determine these initial trim points comparable to flight conditions. To solve this question, a hybrid genetic algorithm is presented in this paper, that combines the quick convergence ability of the quasi-Newton method and the advantages of genetic algorithm, such as global convergence. The trim control vector and the constraint conditions were established in the coordinated-turn based on the helicopter flight dynamic model. The coordinated turn flight of a UH-60 A helicopter was taken as an example to simulate on the experimental platform. Comparisons were made between the trim results and flight test data and there is a good agreement among them, and the efficiency of the algorithm presented is verified. It is a general method that can be applied to trim the helicopter of different flight conditions.

Modeling and control for ultra-low altitude cargo airdrop

PurposeThe purpose of this paper is to model the aircraft-cargo’s coupling dynamics during ultra-low altitude heavy cargo airdrop and to design the aircraft’s robust flight control law counteracting its aerodynamic coefficients perturbation induced by ground effect and the disturbance from the sliding cargo inside.Design/methodology/approachAircraft-cargo system coupling dynamics model in vertical plane is derived using the Kane method. Trimmed point is calculated when the cargo fixed in the cabin and then the approximate linearized motion equation of the aircraft upon it is derived. The robust stability and robust H∞ optimal disturbance restraint flight control law are designed countering the aircraft’s aerodynamic coefficients perturbation and the disturbance moment, respectively.FindingsNumerical simulation shows the effectiveness of the proposed control law with elevator deflection as a unique control input.Practical implicationsThe model derived and control law designed in the paper can be applied to heavy cargo airdrop integrated design and relevant parameters choice.Originality/valueThe dynamics model derived is closed, namely, the model can be called in numerical simulation free of assuming the values of parachute’s extraction force or cargo’s relative sliding acceleration or velocity as seen in many literatures. The modeling is simplified using Kane method rather than Newton’s laws. The robust control law proposed is effective in guaranteeing the aircraft’s flight stability and disturbance restraint performance in the presence of aerodynamic coefficients perturbation.

National Quality Assurance & Improvement System (NQAIS) – Medicine

Introduction: Length of stay varies by hospital, admission type, diagnosis and clinical team, while waiting lists, emergency department throughput and delayed discharges generate frustration for patients and health professionals. The National Acute Medical Programme (NAMP) was established in 2010 by the Royal College of Physicians of Ireland and the Health Service Executive (HSE) to enable: - Swifter access to senior decision makers - Swifter access to investigations and interventions - Reduced overnight admissions - Shorter average length of stay (AvLOS) - Optimal use of Acute Medical Assessment Unit (AMAU) NQAIS Medicine: NQAIS Medicine supports NAMP and hospitals to re-engineer care delivery based on interdisciplinary co-operation using timely and relevant data. The system was developed by Health Intelligence, Health & Wellbeing, Health Service Executive, in collaboration with NAMP and OpenApp. It is deployed under the governance framework of the Joint NQAIS (Surgery & Medicine) Steering Group. NQAIS Medicine is a web-enabled, role-based quality assurance and improvement system providing an evidence base to inform decisions. Innovative “diamond” plots (simplified box and whisker) and trend plots in easily understood “ribbons” are displayed by hospital (or group), admission category, diagnosis and clinical team. Patient profiles (age, co-morbidity, admission day/time etc.) can be flexibly explored. Selected records can be identified for chart review. The predictive impact of meeting targets are summarised by the novel computation of the potential “number of beds freed per day” which could be made available for other patients. The Clinical Classification System (CCS) - Agency for Healthcare Research & Quality (AHRQ) - converts the principal (admission) diagnosis into clinically meaningful groups. The Charlson Index is used to convert secondary diagnoses into a “complexity” score. AvLOS and readmission rates are dynamically analysed in the context of data-driven targets: AvLOS per CCS diagnosis as achieved by top quartile (Irish) teams; 50%+ all acute medical admissions should be routed via the AMAU of whom 25% should be discharged home. AvLOS is computed for each CCS nationally in the context of its trim point - national 75%ile + 3 X (25% – 75%ile). If the target AvLOS is achieved and AMAU usage optimised, the potential impact is translated into simple summary metrics - beds used within, close to and off target. These metrics identify and prioritise processes in hospital most likely to reveal potential learning and reengineering opportunities for using beds for other patients. Implementation: A train-the-trainers approach within hospital groups, guided by a national training team, supports the national deployment of the system. NQAIS Medicine combines science and art to provide detailed but very understandable feedback to unequivocally identify areas for review. Modest meeting of targets would transform a system struggling to cope with increasing demands. The successful implementation of NAMP, using NQAIS Medicine as a continuous improvement enabler, requires sustained focus and commitment from clinicians and management alike. The system is currently being deployed to all acute hospital in Ireland, and is being extended to encompass surgery and other domains within a single system referred to as NQAIS Clinical.

Small morphing wing aerial vehicle dynamic modelling basing on simulation and flight test

Morphing wings can be more stable and more robust to gust in the flow of low Reynolds number compared with normal fixed wings. In contrast to big airplanes, small ones with morphing wings can rely on passive wing deformation to reduce disturbance rather than active control by multi-control surfaces. Engineering modelling for the prototype in the paper ignores aeroelastic effect because of the small deformation amount in terms of static aeroelasticity and high modal frequency in terms of dynamic aeroelasticity. Aerodynamic coefficients are determined through flight test and software calculation. An accurate trim condition is solved using a 6DOF flight model. The preciseness of the trim result is proved by comparing experiment results. The longitudinal and lateral state-space equations benchmarked against trim point is built. Dynamic response of the non-control perturbation motion are analysed.

Nonlinear trajectory transcriptions for aircraft loss of control recovery

Aircraft loss of control may occur when the interpolated gain set is such that it destabilizes the linear aircraft. That is, when linear aircraft is not asymptotically stable, the nonlinear aircraft loss of control occurs. Even in cases where interpolation stabilizes the linear aircraft, loss of control may also occur when the gains are not reconfigured within the stability region associated with the controller. Hence, understanding nonlinear aircraft navigation and control algorithms using such controllers that originate from various linear design methods becomes a challenging problem. It prompts to simulate the nonlinear aircraft trajectories within the distorted stability regions of the linear controllers interpolated when each one of them is designed to regulate an equilibrium (or trim) point. In this article, the gain interpolation parameter that would lead to loss of control is presented. Then, techniques to avert the loss of control using nonlinear trajectory transcriptions are presented. A 3-degree-of-freedom aircraft model of Langelaan is considered for illustrations.

Fast Computable Recoverable Sets and Their Use for Aircraft Loss-of-Control Handling

The paper describes a framework for constrained flight control based on the use of recoverable sets that are chained together to guide constrained admissible transitions between trim points. The recoverable set is the set of all states for which there exists a control sequence such that the subsequent response is guaranteed to satisfy the imposed constraints. The constraints can reflect actuator range/rate limits, safety limits, as well as ranges of validity of the aircraft model. Because in aircraft loss-of-control situations, fast onboard computations are necessary, the approach to computing recoverable sets in this paper exploits linear discrete-time models (which can be generated via onboard system identification and reflect effects of failures and degradations) and recovery sequences generated, either by a stable linear finite-dimensional subsystem or through a reset of the dynamic controller states or of its set points. With this approach, only subsets of the full recoverable set can be computed, however, these recoverable subsets possess certain desirable control invariance properties, and their computations are simple; moreover, the onboard generation of the recovery sequence reduces to a low-dimensional quadratic programming problem. The applications of this approach to longitudinal and lateral linearized and nonlinear aircraft flight models are reported.

Aero-thermo-dynamic analysis of a low ballistic coefficient deployable capsule in Earth re-entry

The paper deals with a microsatellite and the related deployable recovery capsule. The aero-brake is folded at launch and deployed in space and is able to perform a de-orbiting controlled re-entry. This kind of capsule, with a flexible, high temperature resistant fabric, thanks to its lightness and modulating capability, can be an alternative to the current “conventional” recovery capsules. The present authors already analyzed the trajectory and the aerodynamic behavior of low ballistic coefficient capsules during Earth re-entry and Mars entry. In previous studies, aerodynamic longitudinal stability analysis and evaluation of thermal and aerodynamic loads for a possible suborbital re-entry demonstrator were carried out in both continuum and rarefied regimes. The present study is aimed at providing preliminary information about thermal and aerodynamic loads and longitudinal stability for a similar deployable capsule, as well as information about the electronic composition of the plasma sheet and its possible influence on radio communications at the altitudes where GPS black-out could occur. Since the computer tests were carried out at high altitudes, therefore in rarefied flow fields, use of Direct Simulation Monte Carlo codes was mandatory. The computations involved both global aerodynamic quantities (drag and longitudinal moment coefficients) and local aerodynamic quantities (heat flux and pressure distributions along the capsule surface). The results verified that the capsule at high altitude (150km) is self-stabilizing; it is stable around the nominal attitude or at zero angle of attack and unstable around the reverse attitude or at 180° angle of attack. The analysis also pointed out the presence of extra statically stable equilibrium trim points.

Reliability Assessment of Actuator Architectures for Unmanned Aircraft

A reliability assessment framework is presented for small unmanned aerial vehicles. The analysis considers several candidate architectures with different numbers of controllable surfaces and servos. It is assumed that a servo fault detection algorithm is available and affected by known rates of false alarms and missed detections. The aircraft flight envelope is analyzed to determine the fault levels for which the aircraft can still be flown at a trim point. For these “flyable” fault levels, it is assumed that the flight control law can be reconfigured to safely land the aircraft. Finally, the probability of catastrophic failure is estimated based on the histogram of (pre-fault) control command distributions, mean time between failure of the individual servos, and missed detection and false alarm rates. In applying the framework to assess the reliability of the candidate architectures, several interesting observations on design trade-offs are made.

Aerodynamic design of a re-entry capsule for high-speed manned re-entry

Aerodynamic configuration of the capsule for manned re-entry at the second cosmic speed requires a relatively higher lift-to-drag ratio compared to the configuration for manned re-entry at the first cosmic speed. The present capsule configurations with high lift-to-drag ratio, for instance, the Apollo and the Orion, have secondary statically stable trim points, which is highly undesirable since it would cause disastrous result once the vehicle flies at the secondary trim point during re-entry. In the present study, effective design methodology for improving the monostability characteristics of re-entry capsules without compromising the lift-to-drag ratio is revealed based on flow field characteristics analysis. As a result, aerodynamic configuration design of a capsule with high lift-to-drag ratio and monostable characteristics is proposed. And the aerodynamic characteristics of the capsule ranging from subsonic to hypersonic speed are predicted by numerical simulations.

Robust LPV autopilot design for hypersonic reentry vehicle

Purpose – The purpose of this paper is to design a robust angle-of-attack (AOA) tracking control system for the hypersonic reentry vehicle (HRV) based on the linear parameter varying (LPV) theory, as the aerodynamic coefficients of the hypersonic vehicle vary quickly during the reentry phase. Design/methodology/approach – First, longitudinal moment trim is done along the desired flight trajectory. The linearized system at each trim point is built and the dynamic characteristics analysis is made. Then the LPV control law with parameter-dependent quadratic Lyapunov function (PDQLF-LPV) is applied to design the AOA tracking autopilot at each trim point. Frequency performance of the autopilot is assessed and the step response simulation is conducted to validate the effectiveness of the control system. Finally, actual AOA command tracking simulations based on the time-varying nonlinear model are carried out to test the correctness and robustness of the PDQLF-LPV autopilot. Findings – Analysis results demonstra...

LieTrICP: An improvement of trimmed iterative closest point algorithm

We propose a robust registration method for two point sets using Lie group parametrization. Our algorithm is termed as LieTrICP, as it combines the advantages of the Trimmed Iterative Closest Point (TrICP) algorithm and Lie group representation. Given two low overlapped point sets, we first find the correspondence for every point, then select the overlapped point pairs, and use Lie group representation to estimate the geometric transformation from the selected point pairs. These three steps are conducted iteratively to obtain the optimal transformation. The novelties of this algorithm are twofold: (1) it generalizes the TrICP to the anisotropic case; and (2) it gives a unified Lie group framework for point set registration, which can be extended to more complicated transformations and high dimensional problems. We conduct extensive experiments to demonstrate that our algorithm is more accurate and robust than several other algorithms in a variety of situations, including missing points, perturbations and outliers.

Affine LPV-modeling for the ADDSAFE benchmark

General approaches are presented for approximating the nonlinear dynamics of a commercial aircraft model by affine linear parameter varying (LPV) models that are used for the design of fault detection and diagnosis (FDD) systems. One part of the paper deals with the approximation of the nonlinear actuator dynamics, where a simple analytic model is derived that accurately describes a given nonlinear, (partly) black-box model in the whole flight envelope. The second part presents a new two step approach for generating an affine LPV model for the nonlinear aircraft dynamics. Starting from a given trim-point in the flight envelope, the goal of the first step is to maximize the size of the region around this trim point, for which the simple affine LPV description is still valid. The second step then tries to further simplify the affine LPV model, which helps to reduce both the computational burden for FDD synthesis methods and the order of the linear fractional representations (LFRs) that are generated from these LPV models.