Aerodynamic Data Research Articles

High resolution size characterization of particulate contaminants for radioactive metal waste treatment

To regulate the safety protocols in nuclear facilities, radioactive aerosols have been extensively researched to understand their health impacts. However, most measured particle-size distributions remain at low resolutions, with the particle sizes ranging from nanometer to micrometer. This study combines the high-resolution detection of 500 size classes, ranging from 6 nm to 10 μm, for aerodynamic diameter distributions, with a regional lung deposition calculation. We applied the new approach to characterize particle-size distributions of aerosols generated during the plasma arc cutting of simulated non-radioactive steel alloy wastes. The high-resolution measured data were used to calculate the deposition ratios of the aerosols in different lung regions. The deposition ratios in the alveolar sacs contained the dominant particle sizes ranging from 0.01 to 0.1 μm. We determined the distribution of various metals using different vapor pressures of the alloying components and analyzed the uncertainties of lung deposition calculations using the low-resolution aerodynamic diameter data simultaneously. In high-resolution data, the changes in aerosols that can penetrate the blood system were better captured, correcting their potential risks by a maximum of 42%. The combined calculations can aid the enhancement of high-resolution measuring equipment to effectively manage radiation safety in nuclear facilities.

Flight Performance of Crew Escape System during Pad Abort Condition

As a prologue to the Gaganyaan project, ISRO successfully test fired one of the critical and essential technologies for human space flight, i.e. Crew Escape System (CES) Pad Abort Test on 5 July 2018. This is a strategically important flight as a part of various qualification tests that ensures emergency escape measure to quickly pull the Crew Module (CM) along with the astronauts to a safe distance from the launch vehicle in case of any eventuality. CES along with CM weighing of 12.6 tonnes lifted-off at 07.00 h from the Sounding Rocket Complex Launch Pad, Sriharikota. The vehicle was taken to an altitude of 3 km using a specially developed solid motor with multiple reverse flow and scarfed nozzles with 10 g acceleration. Around 330 sensors on-board were used to measure its performance. The data obtained ensured that the adopted design philosophy and approaches such as CM reorientation, aerodynamic data with jet-on environment, mission sequences, performance of special solid motors, deceleration system, etc. were in order. This provides more confidence to progress to the next level of demonstration as a part of Gaganyaan.

Reduction in aerodynamic drag acting on a commercial vehicle by using a dimpled surface

Advancements in the automotive industry have led to an increasing need to reduce the consumption of fuel, commercial vehicles especially public transport. However, when it comes to commercial vehicles like buses, the reduced mileage stands as a disadvantage. This work proposes a numerical model of a bus (commercial vehicle) with design modifications such as dimples on the surface using the Computational Fluid Dynamics (CFD) to increase the fuel economy. It will help develop the flow structure around the bus to understand the amount of drag. Following the preparation of a generic model of the bus using SOLIDWORKS, the drag coefficient was determined employing a wind tunnel test. The aerodynamic data and the detailed flow structure were studied and analyzed. Accordingly, a modified model was designed, comprising dimpled structures, and the tests repeated. Parameters like dimple position, number of dimples, and dimple orientation were closely observed to understand the performance improvement. It can be an effective method to reduce drag, paving the way to better aerodynamics and increasing fuel economy for such heavy commercial vehicles.

Preliminary Design Method and Prototype Testing of a Novel Rotors Retractable Hybrid VTOL UAV

Conventional hybrid VTOL (Vertical Take-Off and Landing) aircraft are always designed by directly mounting multicopter rotors to a fixed-wing plane. This type of VTOL aircraft may suffer from terrible drag due to the copter rotors. To tackle this problem, this paper describes a novel rotors retractable hybrid VTOL UAV (RR VTOL) that combines the hovering capability of multicopters and the high cruise velocity property of pure FW aircraft. Computational fluid dynamic (CFD) analysis on rotors folding and unfolding models proves that the retractable rotors can increase the lift-to-drag ratio by approximately 2 times compared to that common hybrid VTOL UAV. Moreover, some aerodynamic performance data of the rotors folded model are also acquired. A preliminary design method of this creative UAV is proposed with some new empirical equations derived from many hybrid VTOL UAVs’ data. Emphasis is given on the method for determining the center of gravity position and horizontal tail dimensions as they are the unique aspects of the RR VTOL configuration. Subsequently, the manufacturing process and rotor retraction mechanism of two UAV generations are reviewed and a rapid prototyping method for RR VTOL is summarized. Finally, a successful flight test confirms that the proposed VTOL concept is practical and the drag reduction effect of rotors retracting function is remarkable.

Improvement of Electric Aircraft Endurance through Propeller Optimization via BEM-CFD Methodology

An exhaustive optimization method is developed to minimize the power consumption for a propeller-driven electrical aircraft. The method finds the optimal value for a wide range of geometrical and operational parameters for a target thrust and airspeed. The optimization routine employs BEM for propeller predictions fed with aerodynamic airfoil data obtained from a proposed combined CFD-Montgomerie method which is also validated, furthermore several corrections to account for compressibility, three dimensional, viscous and Reynolds number effects are implemented. This BEM model showed an adequate fitting with experimental data. Additionally, Goldstein optimization via Vortex Theory is employed to design pitch and chord distributions minimizing the induced losses of the propeller. The optimization algorithm is validated through a study case where an existent optimization problem is approached leading to very similar results. Some trends and insights are obtained and discussed from the study case regarding the design of an optimal propulsion system. Finally, CFD simulations of the study case are carried out showing a slight relative error of BEM.

An Investigation of the Longitudinal Dynamic Stability of a Box-Wing Aircraft

This study was undertaken to identify the longitudinal dynamic characteristics of the Box-Wing Aircraft. Research on the Box Wing aircraft gained traction following increasing demand for an environmental friendly aircraft that would reduce noise and pollutant emissions. The Box-Wing Aircraft is renowned for its potential aerodynamics advantages over conventional aircraft. However, due to its departure from conventional configuration, the stability dynamics of the Box Wing has not been fully explored. This study investigates the longitudinal dynamic stability characteristics of a Box Wing Aircraft at the conceptual design level using the J2 Aircraft Dynamics Software. The study integrates empirical mass estimation methods and aerodynamic data generated from a vortex lattice tool into J2 Aircraft Dynamics Software Suite. This was done to assess the Short Period Oscillation and Phugoid responses of a Box Wing Aircraft and a conventional cantilever wing aircraft. For a 2s step input, the short period oscillation induced on the Box Wing was damped out after 19s resulting in damping ratio of 0.68 and the undamped natural frequency is 1.8 <i>rad/s</i>. In the Phugoid mode, the Box Wing Aircraft gives a damping ratio of 0.006 and undamped natural frequency of 0.07 <i>rad/s.</i> On the other hand, the SPO induced on the reference conventional aircraft settles after 14s producing a damping ratio of 0.76 and undamped natural frequency of 3.12 <i>rad/s</i>. The damping ratio and undamped natural frequency for the Phugoid mode of the reference conventional aircraft were however 0.001 and 0.07 <i>rad/s</i> respectively. These results indicate that the reference conventional cantilever wing aircraft possesses acceptable longitudinal dynamic stability characteristics while the Box Wing Aircraft would require stability augmentation systems to improve its flying and handling qualities.

Fluid-Structure Interaction Simulation and Accurate Dynamic Modeling of Parachute Warhead System Based on Impact Point Prediction

To predict a parachute-warhead system's dynamic characteristics and impact point, numerical methods are used to comprehensively predict the large deformations of the parachute during the opening process and the impact point of the system in the terminal landing phase. Fluid-structure interaction simulations based on the arbitrary Lagrangian-Eulerian method are used to study the Disk-Gap-Band parachute's inflation behavior and provide the parachute's aerodynamic parameters at steady state. Based on the obtained aerodynamic data, a nine-degree-of-freedom dynamic model of the parachute-warhead system was established, which was used to predict the landing area of the system by calculating the falling trajectory. Based on the established model, an online impact point prediction program was developed. Finally, the effectiveness and accuracy of the methods were verified by airdrop experiments. The results showed that the methods for the parachute-warhead system modeling during the inflation and terminal descent phases could effectively predict its dynamic characteristics, which could be further applied for precision airdrop missions.

Computational Aerodynamics Study of Competing Conceptual Designs for Advanced Tactical Fighter Aircraft

Advanced tactical fighter (ATF) configurations are bound to perform high angle of attack (AoA) maneuvers. However, existing conceptual design tools available in aerospace industry are based on empirical or potential flows that cannot predict aerodynamic data in nonlinear regimes. High-fidelity computational fluid dynamics algorithms have to be incorporated during conceptual design phase for better assessment between competing configurations. In this research, steady state aerodynamic analysis is conducted to compare four conceptual designs of advanced tactical fighters through Reynolds-averaged Navier–Stokes (RANS) simulations. Prior to the study, two validation test cases were conducted based on ONERA M6 Wing and benchmark unmanned combat air vehicle (UCAV) design to assess the computational setup for the problem. Pressure based solver is used to model the flow field in subsonic, transonic and supersonic regimes at sea level for all four competing designs. The quantitative results include the aerodynamic forces and the longitudinal stability coefficient comparisons among the models and its components. The qualitative analyses include pressure distribution, eddy shedding and behavior of vortices at varying flow angle. Additionally, the empirical estimation for interpolation and post-stall extrapolation are carried out for further flight performance studies.

Development and preliminary assessment of the open-source CFD toolkit SU2 for rotorcraft flows

Computational aerodynamic analyses of rotorcraft main rotor blades are performed in both hover and forward flight. The open-source SU2 code is used for rotor performance prediction. The core of the code is the set of RANS equations, which are solved for determining the flow. In hover, both steady-state and time-accurate modelling techniques of varying complexity are used and assessed. Simulation specific parameters which have a significant influence on the solution are also addressed. In forward flight, the code is developed to include the main rotor blade kinematics which is a prerequisite for modelling a trimmed rotor. Two databases are used for the validation of the rotor performance prediction. The renowned Caradonna–Tung experimental tests of a model rotor are used to evaluate the pressure distribution along the blade during hover. The extensive aerodynamic and aeroacoustic data survey of the AH-1G Cobra helicopter is used to assess the pressure distribution at different advancing and retreating azimuth angles during forward flight. The prediction capabilities of the solver in terms of rotor performance are demonstrated and are overall in good agreement with the measured data.

Comparisons of Aerodynamic Data with the Main Wind Force–Resisting System Provisions of ASCE 7-16. I: Low-Rise Buildings

AbstractThis study assesses the wind load provisions for low-rise buildings with Chapters 27 and 28 in ASCE 7-16 and analyzes the dependence of the aerodynamics on the building geometric parameters...

Comparisons of Aerodynamic Data with the Main Wind Force–Resisting System Provisions of ASCE 7-16. II: Mid- and High-Rise Buildings

AbstractThe purpose of this study is to investigate the effects of building geometry on tall-building aerodynamics, accompanying the analysis for low-rise buildings found in the companion paper. Ae...

FADS based aerodynamic parameters estimation for mars entry considering fault detection and tolerance

Flush air data system (FADS) is able to sense pressure distribution on the surface of the vehicle's forebody during Mars entry phase. Existing studies on the FADS-based aerodynamics data estimation for Mars entry have not attached adequate attention to the potential FADS fault, which should not be ignored in practice from a reliability viewpoint. In this paper, a hybrid method is proposed to simultaneously achieve aerodynamic parameters estimation of Mars entry vehicle and FADS fault detection and tolerance. The estimation procedure merges triples algorithm based coarse estimation with least-square based precise estimation. Meanwhile, a fault identifier matrix is embedded in the estimation procedure, which is worked out from chi-square detection algorithm. The estimation with FADS fault tolerance is realized by the fusion of chi-square detection, hybrid estimation, and random sample consensus algorithms. Comparison simulations demonstrate the effectiveness of the proposed technique.

An estimation method for the fuel burn and other performance characteristics of civil transport aircraft during cruise: part 2, determining the aircraft’s characteristic parameters

ABSTRACTA simple yet physically comprehensive and accurate method for the estimation of the cruise fuel burn rate of turbofan powered transport aircraft operating in a general atmosphere was developed in part 1. The method is built on previously published work showing that suitable normalisation reduces the governing relations to a set of near-universal curves. However, to apply the method to a specific aircraft, values must be assigned to six independent parameters and the more accurate these values are the more accurate the estimates will be. Unfortunately, some of these parameters rarely appear in the public domain. Consequently, a scheme for their estimation is developed herein using basic aerodynamic theory and data correlations. In addition, the basic method is extended to provide estimates for cruise lift-to-drag ratio, engine thrust and engine overall efficiency. This step requires the introduction of two more independent parameters, increasing the total number from six to eight. An error estimate and sensitivity analysis indicates that, in the aircraft’s normal operating range and using the present results, estimates of fuel burn rate are expected to be in error by no more than 5% in the majority of cases. Initial estimates of the characteristic parameters have been generated for 53 aircraft types and engine combinations and a table is provided.

Research on Calculation Method for Acoustic Scattering of Helicopter Noise

A new calculation method of helicopter rotor/fuselage acoustic scattering is developed. Firstly, a CFD analysis model is developed to simulate flow field of the rotor, which is based on the motional embedded grid system and RANS equations, and provides aerodynamic data for rotor noise calculation. Then, FW-H equations are employed to calculate the aeroacoustic characteristics of isolated rotor, and G 1A formulas are applied to calculate the rotor acoustic gradient to provide boundary condition for acoustic scattering. Based on these, the time-domain equivalent source method is applied to calculate acoustic scatter field, and the total acoustic field that considered the fuselage scatter is superposed by isolated rotor acoustics and the scatter one. Finally, the numerical simulations of helicopter main-rotor/fuselage and tail-rotor/fuselage scatter effect are conducted by using the developed models. The results indicate that the helicopter fuselage has important scatter effect on the high frequency acoustics of main rotor and tail rotor, and the acoustic scatter effect become more obvious with the smaller space between the main rotor (tail rotor) and fuselage.

Aerodynamic Similarity of Winged and Lifting Space Vehicles

The aerodynamic data of winged and lifting space vehicles are considered. Recently, there was an investigation with respect to the longitudinal aerodynamics of these vehicles in the supersonic and hypersonic Mach number regimes with the astonishing result that for the lift and drag coefficients a similarity with regard to the geometrical configuration could be observed. It was shown that by introducing a homogenizing reference area with which the related quantities were made dimensionless, fully similar solutions of the lift coefficient and partly similar solutions of the drag coefficients are achieved. In the present paper this investigation is extended to the pitching moment coefficient. Further a generalized fit equation on the basis of exponential functions is constructed, which is able to approximate all the longitudinal aerodynamic coefficients for the supersonic and hypersonic Mach numbers. Finally the potential of these findings are demonstrated by some examples.

Using additive manufactured parametric models for wind tunnel test-based aerodynamic shape optimization

PurposeThe purpose of this paper is to propose a new approach of using additively manufactured parametric models in the wind tunnel test-based aerodynamic shape optimization (ASO) framework and to present its applicability test results obtained from a realistic aircraft design problem.Design/methodology/approachFor aircraft shape optimization, the following three methodologies were used. First, as a validation study, the possibility of using rapid prototyping (RP) model in the wind tunnel test was verified. Second, through the wind tunnel test-based ASO, the application and feasibility of the real fighter aircraft shape optimization were verified. A generic fighter configuration is parameterized to generate various test models using additive manufacturing. Wind tunnel tests are conducted to measure their stability criteria in high angle of attack (AOA). Finally, a computational fluid dynamics (CFD) study was performed and analysis procedures, costs and results compared to the wind tunnel test were compared and reviewed.FindingsRP technology can significantly reduce the time and cost of generating parametric wind tunnel models and can open up new possibilities for wind tunnel tests to be used in the rigorous aerodynamic design loop. There was a slight difference between the results of the RP model and the metallic model because of rigidity and surface roughness. However, the tendency of the aerodynamic characteristics was very similarly predictable. Although there are limitations to obtaining precise aerodynamic data, it is a suitable method to be applied to comparative studies on various shapes with large geo-metric changes in the early phase of design. The CFD analysis indicates that the wind tunnel-based ASO using the RP model shows the efficiency corresponding to the CFD shape optimization.Research limitations/implicationsThe RP parametric models may have various assembly error sources and rigidity problems. The proposed methodology may not be suitable for collecting the accurate aerodynamic database of a final design; rather, the methodology is more suitable to screen out many configurations having fairly large shape variation in the early stage of the design process.Practical implicationsThe wind tunnel test-based ASO can replace or supplement CFD-based ASO. In areas where CFD accuracy is low, such as high AOA flight characteristics, RP model wind tunnel-based ASO can be a research method that can secure both efficiency and accuracy advantages, providing ten times more effective in terms of cost and time. The wind tunnel test is used to obtain aerodynamic data at the final stage of shape design. It can be extended to a comparative study of several shapes in the early design phase. This procedure can be applied for both industrial level and educational aircraft design activities.Originality/valueThis study is the application to be applied as a parametric study on the whole aircraft, rather than using the RP model applying a simple partial control surface or configuration change of a part of the wing. The possibility of using the RP model was confirmed by comparing and verifying each other in a medium-sized wind tunnel using a relatively large RP model and a metallic model. It was verified that it can be applied in the shape design process, not the shape verification in the traditional design procedure, and a comparison with the CFD method was also performed. With further development and validation efforts, the new design framework may become an industrial standard for future aircraft development.

Using a Proper Orthogonal Decomposition representation of the aerodynamic forces for stochastic buffeting prediction

High-performance aircraft often suffer from the consequences of tail buffeting at moderate subsonic Mach numbers and medium to high angles of attack. The impact of the aircraft’s highly unsteady flow field on the tails can result in significant structural fatigue and degraded handling qualities. Various methods have been developed to predict tail buffeting. Stochastic response methods are among frequently used approaches. For such methods the size of the excitation data set can become an issue, especially when the auto- and cross-spectra of all available excitation signals on the configuration are considered. The present paper demonstrates how to modify stochastic tail buffeting prediction methods using Proper Orthogonal Decomposition (POD). The approach is based on the modal decomposition of the aerodynamic buffet excitation data set. It notably reduces the computational effort for structural response and loads prediction with limited losses in accuracy while using all power- and cross-spectra of the reduced dataset. The method was applied to the computational buffeting prediction for a generic configuration with double-delta wing and horizontal tail plane (HTP) over a wide range of angles of attack. It was shown that the POD-modes of the aerodynamic buffet excitation resembled the characteristics of configuration’s complex vortical flow field. The predicted structural response and loads converged well with increasing number of POD-modes. With the presented approach, the computational effort of stochastic tail buffeting prediction has been reduced by orders of magnitude compared to the case with the full aerodynamic buffet excitation data set.

Assessment of Impact of Aerodynamic Loads on the Stability and Control of the Gyrocopter Model

Aerodynamic modelling currently relates to development of mathematical models to describe the aerodynamic forces and moments acting on the aircraft. It is a challenging part of aerodynamics that defines a comprehensive approach to using traditional methods and modern techniques to obtain relevant data. The most complicated task for the aerodynamics and flight dynamics is definition, computation and quantification of the aerodynamic description of an object. This paper presents how to determine the aerodynamic load on a gyrocopter and defines the effect on its stability and control. The first step to solution is to develop simpler approximate aerodynamic model - a model that can be used in analysis of aerodynamic load and can represent the aerodynamic properties of the gyrocopter with an acceptable degree of accuracy. Control and stability are very important parts of aircraft characteristics and therefore those characteristics were analyzed in simulation. Finally, the aerodynamic data outputs are assessed in terms of impact of aerodynamic loads on stability and control of the gyrocopter model.

An improved aeroservoelastic modeling approach for state-space gust analysis

As is well known, gust responses of an elastic aircraft can be predicted by both time-domain and frequency-domain methods. The frequency-domain method can produce accurate predictions as long as the obtained aerodynamic data in frequency-domain is accurate. The time-domain method might involve the problem of degraded accuracy due to the modeling errors resulting from the rational function approximation (RFA) of the frequency-domain aerodynamic forces. However, the state-space time-domain model can efficiently deal with a large number of calculations needed for determining the worst-case gust loads, and is very convenient for the design of an aeroelastic controller. In view of the disadvantages of traditional RFA, this paper proposes an aerodynamic subsystem realization technique, which is used to build time-domain aeroservoelastic state-space model with a high precision. Firstly, the Loewner framework is introduced to generate an accurate descriptor system using frequency-domain aerodynamic data. Then, model decomposition and reduction are applied to the descriptor system in order to generate a reduced order system with guaranteed stability. Finally, the aerodynamic subsystem is expressed in state-space form after extracting the additional direct term. The proposed modeling method of the aerodynamic subsystem does not need iterative calculations and artificial selection of aerodynamic poles, and has the advantages in accuracy, efficiency and numerical stability. The proposed method is verified by an example of a transport aircraft flying into the fields of discrete and continuous gusts. The simulation results demonstrate that the developed state-space aerodynamic model is sufficiently accurate, and the time-domain method established in this paper can achieve the same accuracy as the frequency-domain method.

An aerodynamic variable-fidelity modelling framework for a low-observable UCAV

The extensive amount of aerodynamic data needed for the creation of a Stability & Control data set requires efficient and reliable tools. For this reason, the task of producing a large and dense data set in general relies on ground-facility testing and is conducted after the product definition phase, where, nevertheless, already plenty of aerodynamic data is generated by use of simple low-fidelity aerodynamic methods. However, these methods are not capable of predicting complex non-linear aerodynamic phenomena, which complicates a reliable assessment of new designs and demands more accurate predictions early in the design process. Variable-fidelity surrogate modelling offers an opportunity to include high-fidelity simulation tools at affordable effort by enriching a low-fidelity data set with just a few selected highly accurate results. In this paper, an automated aerodynamic variable-fidelity modelling process is proposed to create such an aerodynamic data set at product definition phase. It is demonstrated on an industrial-relevant, agile, low-observable unmanned combat aerial vehicle. An adaptive sampling strategy is applied which proves to be both efficient and accurate. This is supplemented by a model selection algorithm. Moreover, a comparison with a single-fidelity surrogate modelling process shows the variable-fidelity modelling strategy to be superior.