Airframe Model Research Articles

Airframe design and CFD analysis of light unmanned reconnaissance aircraft

This work deals with a comparative numerical finite volume method of airflow for a 3D airframe model. The analyzed airframe model allowed 3-vector thrust directing while using the geometry of the inverted delta, which increases lift force area, allows for vertical take-off and landing, as well as full control of its flight using rotating propulsion nacelles and an airframe control system. The airframe model dimensions were 3900 x 3250 x 915 mm, enabling the installation of components of specific dimensions. CFD analysis using a four-equation transition SST model with automatic surface roughness and initial conditions: T = 288K, viscosity η= 1.7894∙ 10−5 and was carried out in the FLUENT module of ANSYS 16.0. For each simulation, 2000 iterations were performed with the intention of achieving a convergence of all residuals below 1∙10-3 using a calculation mesh consisting of 500,000 elements and 150,000 nodes. The model uses aerodynamic profile geometries: CLARK Y, NACA 6409, and NACA 2414, which are provided to obtain the optimal aerodynamic characteristics and determine the optimal range of angles of attack for the designed solution. Those angles of attack were in the range of α = 5 ÷ 10° in which the highest excellence coefficient of the carrier airfoil profile occurs. For those values of parameters optimal lift and drag forces were determined

A semi-active shunted piezoelectric tuned-mass-damper for multi-modal vibration control of large flexible structures

A semi-active electromechanical Tuned Mass Damper (SATMD), which consists of a combined spring-piezoelectric device connected to an external resistive-inductive electric circuit is presented, aiming to provide robust multi-modal vibration suppression capabilities in large flexible structures. The combined spring-piezoelectric device provides the stiffness for the resonant mass, and also electromechanical energy conversion and coupling. A reduced-ordered mode superposition structural dynamics model coupled with the SATMD device and the RL shunt electric circuit is formulated, to simulate the response of the integrated structural system. The broadband vibration reduction capabilities of the proposed SATMD are numerically and experimentally evaluated and quantified on a down-scaled simplified airframe model. Very good correlation between numerical and measured results is obtained. The results illustrate that a SATMD configuration with an auxiliary mass equal to 1% of the structural mass and combined resistive-inductive impedance, yields substantial simultaneous vibration suppression over the range of 3 low-frequency structural modes, and can be easily retuned to accommodate in-flight changes in structural and loading parameters.

Wind Tunnel Experiments on an Aircraft Model Fabricated Using a 3D Printing Technique

Experimental tests regarding the M-346 aircraft model made via 3D printing were carried out in order to obtain numerical data and characteristics in the form of graphs of basic aerodynamic forces and coefficients. The tests were carried out for the left side of the airframe model in a clean configuration, without additional suspension equipment; the flight control surfaces and the aerodynamic brake were in neutral positions. Based on the scan of the base model in 1:48 scale using a Nikon Model Maker MMDx laser scanning head, followed by the generation and optimization of some of the airframe elements in SolidWorks software, a test model ready for printing was prepared. Using the MakerBot Print program, the printing parameters were set, and the process itself was completed using a MakerBot Replicator Z18 3D printer. The next step was manual treatment in order to remove the material excess from the melted thermoplastic material, join the elements and appropriately polish the surface of the tested model in order to obtain the desired quality. The test was carried out using a Gunt HM 170 wind tunnel for fixed airflow velocities at variable angles of attack. On this basis, the numerical values of lift force, Pz, and drag force, Px, were obtained; then, the lift force indices, Cz, and drag force indices, Cx, were computed for the steady states, which were for angle α from −12 to 16°. The use of 3D printing contributed to the generation of geometry, which, for research purposes, was scaled down in order to fully use the available measurement space of the wind tunnel. The final stage of the work was to compare the obtained curves of particular characteristics with the literature data.

축소 기체 구조 모델을 이용한 고속 비행 복합형 무인 회전익기의 능동 진동 제어 시험

Active vibration control for a 25 %-scale model of a high-speed compound unmanned rotorcraft was experimentally studied. The compound rotorcraft uses wings, propellers, and a lift-offset coaxial rotor system with two blades per rotor. A small-scale airframe structure was manufactured using aluminum and steel and appropriately represents the dynamic similarity of the full-scale model. A ground vibration test (GVT) was used in this experimental work. An active vibration control system (AVCS) using two force generators and three accelerometers was applied for the active vibration control of the small-scale airframe model. Three accelerometers that measure acceleration signals in the vertical direction of the airframe are located at specific airframe locations: the remote cockpit unit, wing root, and wing tip. A 2P rotor hub pitch moment that produces a linear force generator was used to excite the airframe structure. The AVCS using the Fx-LMS algorithm was implemented using MATLAB Simulink, dSPACE ControlDesk, and MicrolabBox. This experimental study shows that the vibration responses of the airframe in the vertical direction at the specific airframe positions are reduced by 47.54 % ~ 84.67 %, and the test results compare reasonably with the simulation results.

Landing Gear Mechanical Network Synthesis for Improving Comfort at Landing Considering Aircraft Flexibility

The landing impact that occurs during aircraft touchdown results in the development of significant loads and accelerations within the airframe. Accurate knowledge of the landing loads is necessary not only for predicting the resulting stresses for design of the airframe, but also for designing strategies to mitigate the vibratory loads and improve the ride quality. Perceived passenger comfort is dependent both on the magnitude of the acceleration experienced by the passengers and on the frequency content of the vibrations. Using a flexible airframe model to better capture the loading regime and frequency response at landing impact, this study optimizes various single-port (two-terminal) passive mechanical networks that consist of an arrangement of springs, dampers, and inerters to minimize passenger discomfort and peak forces applied to the aircraft. The performance is compared with a baseline oleo-pneumatic shock absorber. First, the candidate layouts are optimized, and then the observations from this exercise are used to synthesize a mechanical network with a desired frequency response. All considered mechanical networks demonstrated the ability to control the frequency content of the input loading, thus resulting in a reduction in accelerations and an improvement in all comfort parameters used in this study over the oleo-pneumatic baseline.

Development of Approximate Prediction Model for 3-DOF Helicopter and Benchmarking with Existing Controllers

<p>Recent trend of living is getting modernized rapidly by the involvement of automatic systems. Within the aviation industry, automatic systems had become heavily reliable by the end of the nineteen centuries. The systems usually require controllable devices with desired control algorithm known as controller. Controllers can be replaced with, almost every mechanical automation aspect where, safety is a serious issue. But it is not easy to adapt a controller with a specific model at the beginning. It is important to predict the model before a controller works on the model and the controller parameters need to be adapted to get maximum efficiency. A 3-DOF (Three Degrees of Freedom) airframe model is an advanced benchmark model of real 3-DOF helicopter. It has the same uncommon model dynamics with nonlinearities, strong duel motor cross coupling system, uncertain characteristics, disturbances dependent, unmodeled dynamics and many more. The 3-DOF airframe model is a well-known platform for controller performance benchmarking. This research paper shows the development of an approximate prediction model of a Three Degrees of Freedom helicopter model and uses the proposed approximate model to observe the performance of an existent hybrid controller. The hybrid controller is the combination of two different controllers named Quantitative Feedback Theory (QFT) controller and Adaptive controller. To achieve the research objective, the proposed mathematical model of this airframe was used to develop transfer function and simulate with the hybrid controller in MATLAB. The performance of the controller based on the proposed heliframe model of 3-DOF helicopter have also been reported added within this paper.</p>

Gain-Scheduled Autopilot Design and Validation for an Experimental Guided Projectile Prototype

This paper investigates the design and validation of a gain-scheduled skid-to-turn autopilot for an experimental canard-guided, fin-stabilized 80-mm-guided projectile prototype in a hardware-in-the-loop (HIL) wind tunnel-based setup. The projectile dynamics are a priori unknown and are determined from open-loop system identification experiments conducted on the test setup. To this end, a nonlinear airframe model is discussed and considered for linearization around equilibrium operating points. The linear model parameters are estimated from several data collection experiments, resulting in a family of linear models. This knowledge is leveraged in a second step to compute a family of robustly stable multiobjective disturbance-rejection and reference-tracking controllers, which are implemented on the HIL test setup using a zero-pole gain interpolation scheme. The time-domain performance of the autopilot is finally assessed by the means of numerical simulations and experimental results gathered from the test setup, demonstrating the excellent suitability of the proposed approach.

Propulsion Airframe Aeroacoustic Integration Effects for a Hybrid Wing Body Aircraft Configuration

An extensive experimental investigation was performed to study the propulsion airframe aeroacoustic effects of a high bypass ratio engine for a hybrid wing body aircraft configuration where the engine is installed above the wing. The objective was to provide an understanding of the jet noise shielding effectiveness as a function of engine gas condition and location as well as nozzle configuration. A 4.7% scale nozzle of a bypass ratio seven engine was run at characteristic cycle points under static and forward flight conditions. The effect of the pylon and its orientation on jet noise was also studied as a function of bypass ratio and cycle condition. The addition of a pylon yielded significant spectral changes lowering jet noise by up to 4 dB at high polar angles and increasing it by 2 to 3 dB at forward angles. In order to assess jet noise shielding, a planform representation of the airframe model, also at 4.7% scale was traversed such that the jet nozzle was positioned from downstream of to several diameters upstream of the airframe model trailing edge. Installations at two fan diameters upstream of the wing trailing edge provided only limited shielding in the forward arc at high frequencies for both the axisymmetric and a conventional round nozzle with pylon. This was consistent with phased array measurements suggesting that the high frequency sources are predominantly located near the nozzle exit and, consequently, are amenable to shielding. The mid to low frequency sources were observed further downstream and shielding was insignificant. Chevrons were designed and used to impact the distribution of sources with the more aggressive design showing a significant upstream migration of the sources in the mid frequency range. Furthermore, the chevrons reduced the low frequency source levels and the typical high frequency increase due to the application of chevron nozzles was successfully shielded. The pylon was further modified with a technology that injects air through the shelf of the pylon which was effective in reducing low frequency noise and moving jet noise sources closer to the nozzle exit. In general, shielding effectiveness varied as a function of cycle condition with the cutback condition producing higher shielding compared to sideline power. The configuration with a more strongly immersed chevron and a pylon oriented opposite to the microphones produced the largest reduction in jet noise. In addition to the jet noise source, the shielding of a broadband point noise source was documented with up to 20 dB of noise reduction at directivity angles directly under the shielding surface.

Clustering of Parameter Sensitivities: Examples from a Helicopter Airframe Model Updating Exercise

The need for high fidelity models in the aerospace industry has become ever more important as increasingly stringent requirements on noise and vibration levels, reliability, maintenance costs etc. come into effect. In this paper, the results of a finite element model updating exercise on a Westland Lynx XZ649 helicopter are presented. For large and complex structures, such as a helicopter airframe, the finite element model represents the main tool for obtaining accurate models which could predict the sensitivities of responses to structural changes and optimisation of the vibration levels. In this study, the eigenvalue sensitivities with respect to Young's modulus and mass density are used in a detailed parameterisation of the structure. A new methodology is developed using an unsupervised learning technique based on similarity clustering of the columns of the sensitivity matrix. An assessment of model updating strategies is given and comparative results for the correction of vibration modes are discussed in detail. The role of the clustering technique in updating large-scale models is emphasised.

Clustering of Parameter Sensitivities: Examples from a Helicopter Airframe Model Updating Exercise

The need for high fidelity models in the aerospace industry has become ever more important as increasingly stringent requirements on noise and vibration levels, reliability, maintenance costs etc. come into effect. In this paper, the results of a finite element model updating exercise on a Westland Lynx XZ649 helicopter are presented. For large and complex structures, such as a helicopter airframe, the finite element model represents the main tool for obtaining accurate models which could predict the sensitivities of responses to structural changes and optimisation of the vibration levels. In this study, the eigenvalue sensitivities with respect to Young's modulus and mass density are used in a detailed parameterisation of the structure. A new methodology is developed using an unsupervised learning technique based on similarity clustering of the columns of the sensitivity matrix. An assessment of model updating strategies is given and comparative results for the correction of vibration modes are discussed in detail. The role of the clustering technique in updating large-scale models is emphasised.

From Water Tunnel to Poststall Flight Simulation: The F/A-18 Investigation

Introduction T HE prediction of poststall maneuvers and departure characteristics of high-performance aircraft continues to be challenging,1;2 due, in particular, to the highly three-dimensional nature of the motions and aerodynamic responses. A case in point is the F/A-18 aircraft (Fig. 1), which has experienced departures from controlled  ight, typically involving large sideslip excursions with attendant high body-axes rate variations leading to high angles of attack ® (Ref. 4). The lateral–directional departure susceptibility of the F/A-18 provided part of the motivation within the NASA High-Alpha Technology Program for the improvement of high-alpha predictionsbased on ground testing. The gyrations and responsesexperiencedon departureare intrinsicallynonplanar, that is, involvingarbitraryrotationsnot alignedwith theprincipalaxesof the aircraft.3 When the  ight mechanical deŽ ciencies of an F/A-18 simulation were identiŽ ed, a Canadian programwas undertaken to enhance simulator Ž delity.6 The areas in which the airframe model could be improved included the poststall aerodynamicdatabase, the multivariabletable lookuparchitecture,and the aerodynamicmodel. The present paper focuses on the improvements to the database. The shortcomings of the poststall aerodynamic database were attributed largely to facility interference and scaling effects not accounted for and gaps in the parameter space covered. The original rotary database for F/A-18 was generated in the NASA Langley Research Center Spin Tunnel.11 The F/A-18 poststall maneuvering aerodynamics is dominated by forebody/leading edge extension (LEX) vortex interactions at moderately high angles of attack (30 55 deg). In the light of the sensitivity of the vortex interactions to sideslip angle (Ref. 12), it was deemed necessary to generate a more detailed rotary database to deŽ ne the associated relationshipsover a large sideslip range. When the advantages of the water-tunnel approachwere considered, it seemed appropriate to use this ground-testmethodology to generate supplementary data. The orbital platform rotary balance system (OPLEC) at the Institute for Aerospace Research (IAR) water tunnel, a diagnostic tool for studying support interference7;8 and unsteady wall interference, was subsequently used in highresolution tests as well as diagnostic experiments15i18 on the baseline F/A-18. A thorough understanding of scale effects on vortex

Application of H/sup /spl infin// control to pitch autopilot of missiles

The pitch rate control of a missile is analyzed by an H/sup /spl infin// controller. The system dynamics governing the aerodynamic stability derivatives and airframe structure vibration is formulated in a generalized state space realization form. This formulation facilitates the /spl gamma/-iteration method in solving the associated Ricatti equation of the 4-block H/sup /spl infin// control problem. It is shown that the H/sup /spl infin// controller, because of its loop shaping capabilities, can provide improved tracking performance and stability robustness compared with H/sub 2/ (linear-quadratic-Gaussian (LQG)) and classical proportional-integral-derivative (PID) controller in the rigid airframe model. When airframe flexibility effect is considered, the H/sup /spl infin// controller is also shown to be robust under structure natural frequency variations.

Navier—Stokes Calculations of Rotor—Airframe Interaction in Forward Flight

A method for analyzing the mutual aerodynamic interaction between a rotor and an airframe model has been developed. This technique models the rotor implicitly through the source terms of the momentum equations. A three-dimensional, incompressible, laminar, Navier-Stokes solver in cylindrical coordinates was developed for analyzing the rotor-airframe problem. The calculations are performed on a simplified rotor-airframe model at an advance ratio of 0.1. The airframe surface pressure predictions are found to be in good agreement with wind tunnel test data. Results are also presented for velocity and pressure field distributions in the wake of the rotor.

A robustly stable preliminary control systems design for the YF-16 CCV aircraft

A preliminary control system design for the YF-16 CCV aircraft in its longitudinal mode satisfying its flying qualities specifications is investigated. The design is shown to be robustly stable when the actuator model and the unstable pole in the airframe model of the aircraft are subjected to structured uncertainties. Very recent tools of robust stability analysis are utilized to accomplish this goal.< <ETX xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink">&gt;</ETX>

Investigation of Monte Carlo simulation in FAA program KRASH

In 1971, the U.S. Army first supported the development of computer code KRASH to model the impact dynamics and mechanics of airframes. The Federal Aviation Administration continued this support in 1975. Many enhancements have been added to the initial code, and the current official release version is KRASH 85. The next step in the ongoing advancement of KRASH includes uncertainties in modeling capabilities, which is the contribution of this work. In particular, a Monte Carlo simulation framework has been utilized here to permit the input of parameter uncertainties, and thus allow the output variables to be bound with a degree of statistical confidence. An airframe model was selected and preliminary sensitivity tests were performed on four parameters, specifically the impact surface coefficient of dynamic friction, the internal beam damping constant, the external crushing spring damping ratio, and the material properties, including yield stresses. Results from these preliminary tests showed the model was sensitive to variation in the first three parameters, while it was insensitive to changes in the material properties. Accelerations and impulses were plotted for two of the masses in the model. The means and standard deviations at each time step were calculated and incorporated into the plots. Finally, verification whether the simulation yielded statistically significant results, and confidence bounds for results with large uncertainty are presented. The techniques outlined here are completely extendable to as general a KRASH model as desired.

Automating the development of distributed control software

The Development Framework translates application-specific system specifications into parallel, hard real-time implementations, using methods that are both familiar to developers and optimal for the application. The Development Framework approach applies CASE tools-as well as several new tools-to the development of distributed systems, so designers can concentrate on the control-engineering aspects of their systems. The approach addresses three development phases: specification, software design, and implementation. In the specification phase, the control engineer refines behavioral requirements through simulation and analysis, thereby verifying that the system meets its functional requirements prior to implementation. Once the simulated behavior is satisfactory, the specified behavior is translated into a design. Finally, our tools produce source code, either by automatically generating it or by drawing it from a library. We describe the new and existing tools we apply during each phase. We then demonstrate our approach using an example of a linearized roll-pitch-yaw autopilot and airframe model. >

Comparison of the Extended Kalman Filter and Maximum Likelihood Methods for Parameter Identification of an Autonomously Piloted Vehicle

A six-degree-of-freedom airframe model for an autonomously piloted vehicle (APV) is first constructed containing a number of aerodynamic coefficients. The coefficients model include twenty four parameters. The extended Kaiman filter (EKF) and maximum likelihood (ML) methods are introduced to identifying the aerodynamic coefficients and applied to the APV. The two methods are compared by simulated flight test data. The results show that when initial guesses are poor, the ML method provides a better result. On the contrary, if initial guesses are close to the true values, the iterative EKF method is superior to the ML. Thus, the paper demonstrates that the EKF and/or ML method may be employed to improve the aircraft modeling by processing postflight data.

Correlations of Rotor Wake/Airframe Interaction Measurements with Flow Visualization Data

Narayanan M. Komerath Associate Professor School of Aerospace Engineering Georgia Institute of Technology Atlanta, Georgia 30332 revised version of a paper included in the Proceedings of the 46th Annual Forum of the AHS. ABSTRACT Interaction between the aerodynamics of the rotor and the airframe causes large unsteady pressure fluctuations on rotorcraft airframes. A two-bladed rotor and a hemispherecylinder airframe model are used to study these pressure fluctuations in a wind tunnel, simulating low-speed forward flight conditions. Controlled displacement of the airframe is used to obtain finelyspaced pressure data. These are correlated with azimuth-resolved, quantitative laser sheet visualization of the dynamics of the tip vortices and the inboard vortex sheets from the rotor, along the top of the airframe. Three prominent periodic interactions are observed on the airframe surface. These are the effects of blade motion, the approach and impingement of the tip vortex, and of the vortex sheet. These interactions determine, to a large degree, the overall pressure distribution on the airframe surface, and hence the airframe unsteady airloads. The pressure signatures from the blade passage and the vortex impingement are quite different, and are explained using 2-dimensional models. NOMENCLATURE Cp Pressure coefficient based on freestream dynamic pressure CPmean (Pmean -P∞)/q∞ CPinst (Pinst -P∞)/q∞ CPsteady CP(V∞/Vsteady)2 CPunsteady Punsteady/q∞ CT Rotor thrust coefficient, T/(ρπΩ2R4) h Vertical distance between 2-D vortex core and surface H Vertical spacing between rotor hub center and airframe centerline Pmean Mean static pressure P∞ Freestream static pressure Pinst Instantaneous static pressure Punsteady Unsteady component of static pressure (Pinst Pmean) q∞ Freestream dynamic pressure R Rotor radius T Rotor Thrust U Velocity at surface beneath 2-D vortex Vblade Velocity of blade section Vsteady Far field velocity in a steady reference frame. *: This work was part of the first author's Ph.D. dissertation, performed at the School of Aerospace Engineering, Georgia Institute of Technology. This is a V∞ Tunnel free-stream speed X Streamwise distance from hub center XL Local horizontal coordinate Revised from paper at the AHS Forum, 1990. XN X co-ordinate of airframe nose

Velocity field of a cylinder in the wake of a rotor in forward flight

The rotor and the airframe of a rotorcraft can have severe effects on each other because of flowfield interactions. Data on flowfield characteristics have been obtained using laser velocimetry in an attempt to identify and interpret the features of such interactions. This paper summarizes previous results on the flowfield of the rotor and then describes the effects of the rotor wake on the flowfield around an airframe model. A circular cylinder with a hemispherical nose is used as the airframe model and is instrumented with static pressure taps and flush-mounted microphones. A two-bladed teetering rotor is suspended above it in a wind tunnel. The periodic and time-averaged velocity fields are measured along lines parallel to the airframe axis at a rotor advance ratio of 0.1 and a rotor tip Mach number of 0.29. The dominant features in the velocity distribution over the airframe are seen to be linked to vortex interactions with the airframe surface. The tip vortex and the trailing vortex sheet are seen to be displaced in opposite directions by the airframe. Secondary vortex generation phenomena are observed near the airframe surface during vortex interaction.

Velocity measurements of airframe effects on a rotor in a low-speed forward flight

Aerodynamic interactions between the rotor and the airframe of a rotorcraft can have severe effects and are difficult to predict analytically. To attack this problem, the velocity field of a two-bladed rotor has been measured in a wind tunnel with and without an airframe model in proximity. The periodic and time-averged velocity fields were measured using a laser velocimeter in planes both parallel to and above and below the rotor tip path plane at an advance ratio of 0.1 and a rotor tip Mach number of 0.29. The data were shown to be free of tunnel wall effects. The effect of including the cylindrical airframe model were measured. For the geometry studied, airframe influence on the rotor flowfield was mostly confined to the front half of the rotor disc. Hub effects were noticeable, even with the minimal-sized hub used. Strong vortex interaction effects were observed using strobed laser sheet flow visualization and measured using laser velocimetry.