Trim Conditions Research Articles

Hovering and forward flight of the hawkmoth Manduca sexta: trim search and 6-DOF dynamic stability characterization

We show that the forward flight speed affects the stability characteristics of the longitudinal and lateral dynamics of a flying hawkmoth; dynamic modal structures of both the planes of motion are altered due to variations in the stability derivatives. The forward flight speed ue is changed from 0.00 to 1.00 m s−1 with an increment of 0.25 m s−1. (The equivalent advance ratio is 0.00 to 0.38; the advance ratio is the ratio of the forward flight speed to the average wing tip speed.) As the flight speed increases, for the longitudinal dynamics, an unstable oscillatory mode becomes more unstable. Also, we show that the up/down (wb) dynamics become more significant at a faster flight speed due to the prominent increase in the stability derivative Zu (up/down force due to the forward/backward velocity). For the lateral dynamics, the decrease in the stability derivative Lv (roll moment due to side slip velocity) at a faster flight speed affects a slightly damped stable oscillatory mode, causing it to become more stable; however, the thalf (the time taken to reach half the amplitude) of this slightly damped stable oscillatory mode remains relatively long (∼12T at ue = 1 m s−1; T is wingbeat period) compared to the other modes of motion, meaning that this mode represents the most vulnerable dynamics among the lateral dynamics at all flight speeds. To obtain the stability derivatives, trim conditions for linearization are numerically searched to find the exact trim trajectory and wing kinematics using an algorithm that uses the gradient information of a control effectiveness matrix and fully coupled six-degrees of freedom nonlinear multibody equations of motion. With this algorithm, trim conditions that consider the coupling between the dynamics and aerodynamics can be obtained. The body and wing morphology, and the wing kinematics used in this study are based on actual measurement data from the relevant literature. The aerodynamic model of the flapping wings of a hawkmoth is based on the blade element theory, and the necessary aerodynamic coefficients, including the lift, drag and wing pitching moment, are experimentally obtained from the results of previous work by the authors.

Optimum Design of Compound Helicopters that Use Higher Harmonic Control

The optimal design of a compound helicopter comprised of counterrotating coaxial rotors, a propeller, and optionally a fixed wing is investigated. The blade geometry, azimuthal blade pitch inputs, optimal shaft angle (rotor angle of attack), and division of propulsive and lifting forces among the components that minimize the total power for a given flight condition are determined. The optimal design problem is cast as a variational statement that minimizes the sum of induced and viscous power losses for a prescribed lift, propulsive force, and vehicle trim condition. The rotor, propeller, and wing geometry and control inputs are related to the far-field circulation through a lifting-line model that accounts for experimentally or computationally determined nonlinear lift and drag polars. The variational statement is discretized using a vortex lattice wake, and the resulting nonlinear constrained optimization problem is solved via Newton iteration. Results show that varying the prescribed propulsive force of the system affects the optimal shaft angle and rotor design, that higher harmonic control reduces the total vehicle power loss (inefficiency) in high-speed flight by as much as 15%, and that imposing a maximum allowable lateral lift offset can greatly increase the power loss of the coaxial rotors.

가로축 자동비행시스템 개선에 관한 연구

전투기는 공대지 및 공대공 전투임무를 위한 다양한 외부 무장형상으로 운용된다. 무장발사로 인한 비대칭 무장형상에서는 조종사가 롤트림 입력을 인가하여 수평비행을 유지한다. 그러나 외란이 있는 실제 비행 중에는 수평비행을 위해 정확한 롤트림 입력을 인가할 수 없다. 이러한 상태에서 자동비행시스템 인가 시에는 부정확한 롤트림 명령의 영향으로 롤자세각 오차가 발생하여 롤축 자동비행시스템의 성능을 저하시킨다. 본 논문에서는 이러한 문제를 해결하기 위해 롤축 자동비행시스템 명령 단에 적분기를 설계하여 롤자세각의 정상상태 오차를 제거하였다. 또한, 롤자세각 오차량을 제한하는 제한기를 비행영역에 따라 스케줄링하여 초기 과도응답을 개선하였다. 시뮬레이션 평가한 결과, 설계된 비행제어법칙은 부정확한 롤트림 상태에서 자동비행시스템을 운용 시에 롤자세각 추종 성능을 개선함을 확인할 수 있었다. The fighter aircraft uses several different loading configurations for air-to-surface and air-to-air combat missions. To maintain wings-level flight with an asymmetric weapon configuration, a pilot controls a roll trim system. However, it is difficult to apply an accurate roll trim input for wings-level flight in the actual flight under disturbance. The inaccurate roll trim input degrades the performance of the roll autopilot system. In this paper, to solve this problem, an integrator was additionally designed in the command part of the roll autopilot system. The initial transient response was improved by scheduling the limiter to restrict the roll attitude error. As a result of the evaluation of the simulation for the designed flight control law, the roll attitude following performance was found to be improved in the autopilot system operation under the inaccurate roll trim condition.

Design, development and ground testing of hingeless elevons for MAV using piezoelectric composite actuators

A design methodology is presented to develop the hingeless control surfaces for MAV using adhesively bonded Macro Fiber Composite (MFC) actuators. These actuators have got the capability to deflect the trailing edge surfaces of the wing to attain the required maneuverability, besides achieving the set aerodynamic trim condition. A scheme involving design, analysis, fabrication and testing procedure has been adopted to realize the trailing edge morphing mechanism. The stiffness distribution of the composite MAV wing is tailored such that the induced deflection by piezoelectric actuation is approximately optimized. Through ground testing, the proposed concept has been demonstrated on a typical MAV structure. Electromechanical analysis is performed to evaluate the actuator performance and subsequently aeroelastic and 2D CFD analyses are carried out to see the functional requirements of wing trailing edge surfaces to behave as elevons. Efforts have been made to obtain the performance comparison of conventional control surfaces (elevons) with morphing wing trailing edge surfaces. A significant improvement in lift to drag ratio is noticed with morphed wing configuration in comparison to conventional wing. Further, it has been shown that the morphed wing trailing edge surfaces can be deployed as elevons for aerodynamic trim applications.

CFD를 이용한 컨테이너 선형의 트림별 저항성능 해석

Vessels are traditionally optimized for a single condition, normally the contract speed at the design draft. The actual operating conditions quite often differ significantly. At other speed and draft combinations, adjusting the trim can often be used to reduce the hull resistance. Changing the trim is easily done by shifting ballast water. There are several ways to assess the effect of the trim on the hull resistance and fuel consumption, including in-service measurements, model tests, and CFD. In this paper, CFD is employed for the assessment of the resistance performance according to the trim conditions. The commercial CFD code of the STAR-CCM+ is utilized to evaluate the ship’s resistance performance on a 6,800 TEU container ship. To validate of the effectiveness of STAR-CCM+, the experimental result of the KCS hull form is compared with the result from STAR-CCM+. It is found that the total resistance of the 6,8000 TEU container ship was reduced by 2.6% in the case of a 1-m trim by head at 18knots.

Integrated versus two-loop guidance-autopilot for a dual control missile with high-order aerodynamic model

This paper proposes a novel integrated guidance and autopilot design scheme for endo-atmospheric missiles having forward and rearward aerodynamic controls. The missile aerodynamic, including the aero-coupling effects between the canard and tail sections, is assumed to be non-linear. The value of the integrated design scheme is that it is solved in finite-time horizon to target interception, while maintaining missile angle of attack, canard and tail incidence angles within their aerodynamic limitations. The integrated design includes a feed-forward command generator based on the dual control missile trim conditions. The proposed design scheme is compared against a two-loop approach. The test cases via simulations are promising, achieving reduced control efforts and lower miss distances with the integrated logic.

Memory Die Clustering and Matching for Optimal Voltage Window in Semiconductor

In this paper, we propose a method to optimize the product performance instantly by utilizing the internal voltage trimming circuit for Dynamic Random Access Memory (DRAM) memory. Specifically, we first define the verification wafer as the internal voltage characteristics using the clustering technique. Second, the optimized voltage conditions are applied to a normal wafer being matched with a verification wafer. The proposed method makes the ability to apply a different voltage trimming condition for each dies internal voltage circuit depending on their characteristics, thereby improving the characteristics of the individual dies and reducing the fail bit count (FBC) further. The experimental results on the real-application case show that our proposed method reduces the FBC by 1%-5%, which contributes yield enhancement and quality improvement of DRAM memory by raising the efficiency of the redundancy cell repair in the repair process.

Aerodynamic design optimization of helicopter rotor blades including airfoil shape for forward flight

This study proposes a process to obtain an optimal helicopter rotor blade shape including both planform and airfoil shape for helicopter aerodynamic performance in forward flight. An advanced geometry representation algorithm which uses the Class Function/Shape Function Transformation (CST) is employed to generate airfoil coordinates. With this approach, airfoil shape was considered in terms of design variables. The optimization process was constructed by integrating several programs developed by the author. Airfoil characteristics are automatically generated by an analysis tool where lift, drag, and moment coefficients of airfoil are predicted for subsonic to transonic flow and a wide range of attack angles. The design variables include twist, taper ratio, point of taper initiation, blade root chord, and coefficients of the airfoil distribution function. Aerodynamic constraints consist of limits on power available in hover and forward flight, aerodynamic requirements (lift, drag and moment coefficients) for critical flow condition occurring on rotor blades. The trim condition must be attainable in any flight condition. Objective function is chosen as a combination expression of non-dimensional required power in hover and forward flight.

Estimation of lateral-directional aerodynamic derivatives from flight data using conventional and neural based methods

AbstractThe paper presents the estimation of lateral-directional aerodynamic derivatives (parameters) using conventional and neural based methods from real flight data of Hansa-3 aircraft. The conventional methods such as least squares (LS) and maximum likelihood (ML) require an a priori postulation of mathematical model to estimate the parameters. Whereas the neural-based method such as Neural-Gauss-Newton (NGN) is an algorithm that utilises feed forward neural network and Gauss-Newton optimisation to estimate the parameters and does not require any a priori postulation of mathematical model or solution of equations of motion. In the paper, the LS, ML and NGN methods are applied to lateral-directional flight data in order to estimate parameters. The results obtained in terms of lateral-directional aerodynamic derivatives are reasonably accurate to establish LS, ML and NGN as parameter estimation methods along with NGN method having an additional advantage of non-requirement of a priori mathematical model. The paper also highlights the effect of different types of control inputs on parameter estimation. For this, three types of control inputs were used to generate real flight data. The ailerons and rudder were deflected in the first, the ailerons were deflected while keeping rudder at trim condition in the second and the rudder was deflected while keeping ailerons at trim condition in the third type of control input to generate the real flight data. The paper presents the effect of three different types of control inputs in terms of aerodynamic parameters estimated through conventional and neural based methods using flight data generated through these inputs.

Evolution of attitude control law of an Indian re-entry launch vehicle

Reusable launch vehicle (RLV) is a goal pursued by different aerospace agencies, world-wide. Indian space research organization is also one among them. This paper covers the challenges faced and the solutions developed for control law design of a typical re-entry vehicle. The first stage of the test vehicle, used for technology demonstration mission, is a conventional solid rocket motor. The second stage of the test vehicle is a winged body vehicle termed as technology demonstrator vehicle (TDV). This TDV forms the orbiter stage of a conceptual two-stage-to-orbit vehicle. While developing control law for TDV, the experience gained in developing control law for expendable launch vehicles is fully utilized. Due to the presence of winged-second-stage, the solution for attitude control problem naturally inherits the procedure and technology used for designing control law of an aircraft. There are significant differences a re-entry vehicle has in comparison to both aircraft and launch vehicles. Due to these differences, the final control law depends on both aircraft and launch vehicle control law design methodologies. In this paper the authors share the insights gained in control law design of a RLV. The paper gives the reasons for the selection of a combination of co-ordinate systems, for the development of linear plant model of RLV. It emphasizes the difference in the plant modelling with respect to aircraft dynamics in accounting gravitational force in the linear plant model. Requirement of pseudo trim condition for force balance is described in the paper. The paper highlights the commands to be tracked and variables to be fed back during different regimes of re-entry flight. Issues in control law gain design with respect to selected sensors and their placement are elaborated in the paper. Insights on the aspect of trim scheduling, gain scheduling, and finalization of control law structure are also given in the paper.

Coupled aerostructural topology optimization using a level set method for 3D aircraft wings

The purpose of this work is to develop a level set topology optimization method for an unstructured three-dimensional mesh and apply it to wing box design for coupled aerostructural considerations. The paper develops fast marching and upwind schemes suitable for unstructured meshes, which make the level set method robust and efficient. The method is applied to optimize a representative wing box internal structure for the NASA Common Research Model. The objective is to minimize the total compliance of the wing box. The trim condition that aerodynamic lift must balance the total weight of the aircraft is enforced by allowing the root angle of attack to change. The adjoint method is used to obtain the coupled shape sensitivities required to perform aerostructural optimization of the wing box. Optimum solutions for several aerodynamic and body force load cases, as well as a ground load case, are presented.

Damaged Airplane Trajectory Planning Based on Flight Envelope and Motion Primitives

This paper presents an efficient approach for safe landing trajectory generation of an airplane with structural damage to its wing flying in proximity to local terrain. A damaged airplane maneuvering flight envelope is estimated by analyzing the local stability and flying quality at each trim condition. Trim state distance to a flight envelope boundary provides a novel criterion defined as the safety value index to prioritize choice of trim conditions for postdamage flight. A library of trajectory segments including trim states and transition maneuvers between trim conditions in the local neighborhood are generated using a linear quadratic regulator controller. A potential field strategy is used to rapidly define emergency landing trajectories composed as trim state sequences based on criteria including terrain avoidance, safety value index, and safe landing requirements such as touchdown heading and position, airspeed, glide slope, and bank angle. Cost criteria are balanced by adjusting weights. Landing scenarios are presented to demonstrate the capability of the proposed approach to autonomously plan safe landing trajectories.

CFD 저항 해석을 이용한 컨테이너선 최적 트림 추정법에 대한 연구

The main objective of the present study is to elucidate a correlation between ship resistances from computational fluid dynamics (CFD) and brake horse powers (BHP) from towing tank in container carriers. The tests were conducted for a range of combinations of trim conditions and speeds. To achieve this goal, 295 cases of numerical simulation have been performed using Star-CCM+ which had been statistically verified to predict ship resistances (Lee & Lee, 2014). Based on the normal distribution of resistance errors in all cases of the 4 container carriers, the confidence interval of numerical error was estimated as [-2.33%,+2.42%] with 95% confidence. The correlation coefficients between the ship resistances of CFD and the brake horse powers of the experiments were higher than 0.93. As a result, the numerical calculation of ship resistances is able to be utilized in order to provide a quick guidance in selection of the optimum loading condition.

In-flight estimation of center of gravity position using all-accelerometers.

Changing the position of the Center of Gravity (CoG) for an aerial vehicle is a challenging part in navigation, and control of such vehicles. In this paper, an all-accelerometers-based inertial measurement unit is presented, with a proposed method for on-line estimation of the position of the CoG. The accelerometers' readings are used to find and correct the vehicle's angular velocity and acceleration using an Extended Kalman Filter. Next, the accelerometers' readings along with the estimated angular velocity and acceleration are used in an identification scheme to estimate the position of the CoG and the vehicle's linear acceleration. The estimated position of the CoG and motion measurements can then be used to update the control rules to achieve better trim conditions for the air vehicle.

HYDRODYNAMIC PERFORMANCES OF WARPED HARD CHINE DISPLACEMENT HULL FORM

The results of experimental tests relative to resistance, seakeeping and roll decay of a warped hard chine hull form, suitable for large yacht design are reported. To better investigate the effect of hard chine, a round bilge identical model has been tested. Resistance tests have been performed at FV ranging from 0.90 to 4.14, for three different LCG. Results for resistance, sinkage and trim are given. For the most realistic condition dynamic wetted surface is determined separating pressure and whisker spray areas to allow reliable full scale powering predictions. For the same static trim condition, heave, pitch and accelerations have been measured for head sea at FV = 1.35 and 1.89 in regular waves for different /Lratios at constant wave steepness H/ = 1/50. Roll decay at five speeds and roll at zero speed were investigated. The reported data and considerations represent a reference for the implementation of hard chine hull form in the design of displacement and semidisplacement large yachts.

Consistent Structural Linearization in Flexible Aircraft Dynamics with Large Rigid-Body Motion

This paper investigates the linearization, using perturbation methods, of the structural deformations in the nonlinear flight dynamic response of aircraft with slender, flexible wings. The starting point is the coupling of a displacement-based geometrically nonlinear flexible-body dynamics formulation with a three-dimensional unsteady vortex lattice method. This is followed by a linearization of the structural degrees of freedom, which are assumed to be small in a body-fixed reference frame. The translations and rotations of that reference frame and their time derivatives, which describe the vehicle flight dynamics, can still be arbitrarily large. The resulting system preserves all couplings between rigid and elastic motions and can be projected onto a few vibration modes of the unconstrained aircraft with geometrically nonlinear static deflections at a trim condition. Equally, the unsteady aerodynamics can be approximated on a fixed lattice defined by the deformed static geometry. Numerical studies on a representative high-altitude long-endurance aircraft are presented to illustrate the approach. The results show an improvement compared to those obtained using the mean-axes approximation.

Sails trim optimisation using CFD and RBF mesh morphing

The study is focused on the use of mesh morphing to explore different trims of yachts sails. In particular, four trims of the fore and aft sail of a model-scale sailing yacht were modelled leading to 16 configurations in total. Sail pressure distributions were validated with wind-tunnel measurements for all the 16 configurations, and full verification and validation was performed for one of these conditions. The 16 configurations were modelled with two different approaches: generating a new mesh for each trim condition (standard method) and using a morphed version of the baseline condition. This second novel method, based on the use of radial basis functions to morph the mesh, allows the computational time of exploring different geometries with computational fluid dynamics to be significantly decreased. Good agreement is observed between the pressure distributions computed with new meshes and morphed meshes. In order to show an example of trim optimisation, a metamodel approach is defined for the estimation of the response surface using radial basis function interpolation in the parameter space. Thanks to the continuum nature of morphing approach, the optimal trim angles for the given flow condition could be verified using new full computational fluid dynamic simulations. The original full factorial map of 16 points was replaced with a new map of 9 points with an optimal space filling approach to understand the faithfulness of a reduced metamodel. In both cases optimal point is evaluated using a fine design of experiment table built using the metamodel (41 levels for each parameter). The maximum thrust is achieved at the same trim for both metamodels.Proposed method can be easily extended to a wide number of parameters. Such flexibility is demonstrated in the present paper showing the sensitivity of results with respect to apparent wind angle and heeling angle.

Ship Trim Optimization: Assessment of Influence of Trim on Resistance of MOERI Container Ship

Environmental issues and rising fuel prices necessitate better energy efficiency in all sectors. Shipping industry is a stakeholder in environmental issues. Shipping industry is responsible for approximately 3% of global CO2 emissions, 14-15% of global NOX emissions, and 16% of global SOX emissions. Ship trim optimization has gained enormous momentum in recent years being an effective operational measure for better energy efficiency to reduce emissions. Ship trim optimization analysis has traditionally been done through tow-tank testing for a specific hullform. Computational techniques are increasingly popular in ship hydrodynamics applications. The purpose of this study is to present MOERI container ship (KCS) hull trim optimization by employing computational methods. KCS hull total resistances and trim and sinkage computed values, in even keel condition, are compared with experimental values and found in reasonable agreement. The agreement validates that mesh, boundary conditions, and solution techniques are correct. The same mesh, boundary conditions, and solution techniques are used to obtain resistance values in different trim conditions at Fn = 0.2274. Based on attained results, optimum trim is suggested. This research serves as foundation for employing computational techniques for ship trim optimization.

Fault-Tolerant Control of a Unmanned Aerial Vehicle with Partial Wing Loss

Abstract This paper presents the design, implementation and flight testing of a fault-tolerant flight control system for an unmanned aerial vehicle that can accommodate partial wing loss. A dynamic model for the damaged unmanned aerial vehicle (UAV) is based on an asymmetric six degrees-of-freedom equations of motion model. The effect of partial wing loss on the aerodynamic coefficients, centre of mass location and moments of inertia are calculated as a function of percentage wing loss using vortice lattice techniques and computer assisted design software respectively. The model also accounts for changes in the aileron control authority due to partial control surface loss. The equilibrium manifold of the aircraft state is determined as a function of partial wing loss by solving the nonlinear differential equations with a sequential quadratic programming algorithm. A linear parameter-varying model is then obtained by linearising the nonlinear model over the range of equilibrium states. It is found that the trim state of the aircraft changes significantly over the range of percentage wing loss, but that the stability about the trim condition does not change drastically. An acceleration-based flight control system is designed and its robustness is analysed in terms of maximum variations in gain and phase margins. A flight test is performed where 20% of the wing is jettisoned in-flight. The experimental results show that the flight control system is able to stabilise the damaged aircraft and continue with normal waypoint navigation.

<i>H∞</i>-Based LQR Control for Flexible Air-Breathing Hypersonic Vehicles with Pole Placement

Focusing on a nonlinear longitudinal dynamical model for air-breathing hypersonic flight vehicles (AHFV), we propose a state feedback linearized model on a nominal trim condition. To stabilize the flight of an AHFV in the presence of external disturbances, a newHbased Linear Quadratic Regulator (LQR) control law with pole placement is designed. Indexes forHperformance, quadratic performance and pole placement are considered together. As a result, the robustness of system is improved and the AHFV system is effectively stabilized. Numerical simulation shows that the controller can effectively stabilize the AHFV system with some disturbances and assign the poles into a desired region.