Full-scale Flight Research Articles

Scale Effects in Wind Tunnel Testing: Extrapolation of the Drag Coefficient

Subscale wind tunnel data extrapolation is both a necessary and challenging undertaking. Most readily available wind tunnel facilities can achieve Reynolds numbers that are typically an order of magnitude lower than those of the full-scale flight article. Consequently, the wind tunnel data need to be scaled, with the Reynolds number disparity impacting force and moment measurements. Although computational fluid dynamics is ascending as a valuable tool for enabling scaling, the most common extrapolation methodologies are analytic and semi-empirical in nature. Consequently, a new approach is presented to scale both the minimum drag coefficient and the sectional pressure drag coefficient that uses the subscale data to define extrapolation exponents. Implementation requires a minimum of two subscale tests to enable specification of the exponents determined through coalescence of the extrapolated subscale tests. The approach is simple to implement and can be applied to both airfoils and wings. Numerous cases are presented for method validation.

Correlation of High-Speed Tiltrotor Stability Predictions with Test Data and Parametric Study

High-speed stability of tiltrotor was studied. The University of Maryland’s Maryland Tiltrotor Rig (MTR) was chosen for the analysis due to availability of properties and test data, and its interesting high-stability behavior observed in the Glenn L. Martin wind tunnel in August 2022. A Rotorcraft Comprehensive Analysis System (RCAS) model of the MTR gimbaled hub was built in addition to the University of Maryland Advanced Rotorcraft Code-II (UMARC-II) model from previous work. The objective is threefold: i) validate RCAS tiltrotor stability predictions, ii) shed light on the high-stability behavior of the MTR, and iii) find ways to lower the instability speed of the MTR for future wind tunnel tests. Trim collective for freewheeling and stability predictions were compared with wind tunnel test data up to 200 knots. RCAS and UMARC-II predictions showed good agreement with each other and the test data. Predictions show that MTR is stable up to 215 knots (490-knots full-scale flight) although the wing is only 18% thick (current technology is 23%). A parametric study was carried out. The impact of wing stiffness, pitch-flap coupling (δ3 angle), lag stiffness, blade chord, number of blades, pylon mass, pylon center of gravity (c.g.), pylon location, and rotor speed was studied. MTR’s pylon c.g. is unconventionally behind the wing elastic axis. It was found that this significantly improved stability. This behavior is not specific to MTR; full-scale aircraft stability can also be improved by moving the pylon c.g. backward if wing beam is the least stable mode. A combination of forward pylon c.g., reduced rotor speed, and increased blade chord reduced the instability speed by more than 55 knots to near 160 knots, helping researchers obtain high-quality test data in the upcoming Glenn L. Martin wind tunnel tests.

Towards Structural and Aeroelastic Similarity in Scaled Wing Models: Development of an Aeroelastic Optimization Framework

The pursuit of more efficient transport has led engineers to develop a wide variety of aircraft configurations with the aim of reducing fuel consumption and emissions. However, these innovative designs introduce significant aeroelastic couplings that can potentially lead to structural failure. Consequently, aeroelastic analysis and optimization have become an integral part of modern aircraft design. In addition, aeroelastic testing of scaled models is a critical phase in aircraft development, requiring the accurate prediction of aeroelastic behavior during scaled model construction to reduce costs and mitigate the risks associated with full-scale flight testing. Achieving a high degree of similarity between the stiffness, mass distribution and flow field characteristics of scaled models and their full-scale counterparts is of paramount importance. However, achieving similarity is not always straightforward due to the variety of configurations of modern lightweight aircraft, as identical geometry cannot always be directly scaled down. This configuration diversity has a direct impact on the aeroelastic response, necessitating the use of computational aeroelasticity tools and optimization algorithms. This paper presents the development of an aeroelastic scaling framework using multidisciplinary optimization. Specifically, a parametric Finite Element Model (FEM) of the wing is created, incorporating the parameterization of both thickness and geometry, primarily using shell elements. Aerodynamic loads are calculated using the Doublet Lattice Method (DLM) employing twist and camber correction factors, and aeroelastic coupling is established using infinite plate splines. The aeroelastic model is then integrated within an Ant Colony Optimization (ACO) algorithm to achieve static and dynamic similarity between the scaled model and the reference wing. A notable contribution of this work is the incorporation of internal geometry parameterization into the framework, increasing its versatility and effectiveness.

Development of Software for Automatic Determination of Drag on Ballistic Track

The subject of this article is the development of software and algorithmic software for conduct and result processing automation of full-scale ballistic experiments on ballistic routes. At the beginning of the work, attention is focused on the importance of conducting full-scale flight tests of small unguided aircraft during research and design work on their development and modernization. Various methods of measuring velocity in a full-scale or bench-mounted ballistic experiment are considered. The description of the method used in this work based on the determination of the average speed on the site according to the readings of two sensors and the implementation of this method in chronographs of open or closed type is given. Based on this, it follows the presentation of the essence of the conducted ballistic experiments to determine the aerodynamic characteristics of small-sized aircraft, including the longitudinal force aerodynamic coefficient. The transition from the specified coefficient to the coefficient of drag during the experiment on a ballistic track is justified. A description is given of obtaining an analytical expression based on the law of 'living forces' to determine the drag coefficient of an uncontrolled passively flying statically stable object in a ballistic experiment. An algorithm for conducting a ballistic experiment in 'manual' mode is presented, and the automation elements of this algorithm proposed by the authors are presented. As a justification of expediency, a comparative analysis of the 'manual' and automated variants of the experiment was carried out. The procedure for working with the developed software and mathematical software is described. In conclusion, the verification of the obtained results of the ballistic experiment with the results of digital modeling in the developed software and mathematical complex was carried out.

MATHEMATICAL MODEL OF HOT-AIR BALLOON STEADY-STATE VERTICAL FLIGHT PERFORMANCE

Vertical flight performance of Lighter-than-Air free hot-air balloons is derived and discussed. Novel mathematical model using lumped-parameters has been used to model balloon flight dynamics and steady-state performance in particular. Thermal model was not treated as the super-heat is under the control of aeronauts/pilots. Buoyancy or gross lift, net or effective lift, specific lift, and excess specific lift were derived for a general single envelope balloon and can be applied to hot-air, gas and hybrid balloons. Rate-of-climb, absolute ceiling, rate-of-descent, and the maximum rate-of-descent or the uncontrolled terminal descent have all been modeled and sample computations performed for AX8 or AX9 FAI-class hot-air balloons. Lifting index or the specific net/effective lift have been computed treating ambient and hot air as ideal gases at various pressure altitudes and representative envelope temperatures. Drag coefficient in upward and downward vertical flights have been chosen based on best available data. Experimental scale and full-scale flight tests are suggested for more accurate estimates of external aerodynamics in vertical balloon flights. CFD computations of coupled inner- and external-flows are also recommended in future efforts. Knowledge of free balloon’s vertical performance is essential in flight planning and operational safety of flight.

Enhancement of light aircraft 6 DOF simulation using flight test data in longitudinal motion

Abstract This paper proposes a procedure to improve the accuracy of the light aircraft 6 DOF simulation model by implementing model tuning and aerodynamic database correction using flight test data. In this study, the full-scale flight testing of a 2-seater aircraft has been performed in specific longitudinal manoeuver for model enhancement and simulation validation purposes. The baseline simulation model database is constructed using multi-fidelity analysis methods such as wind tunnel (W/T) test, computational fluid dynamic (CFD) and empirical calculation. The enhancement process starts with identifying longitudinal equations of motion for sensitivity analysis, where the effect of crucial parameters is analysed and then adjusted using the model tuning technique. Next, the classical Maximum Likelihood (ML) estimation method is applied to calculate aerodynamic derivatives from flight test data, these parameters are utilised to correct the initial aerodynamic table. A simulation validation process is introduced to evaluate the accuracy of the enhanced 6 DOF simulation model. The presented results demonstrate that the applied enhancement procedure has improved the simulation accuracy in longitudinal motion. The discrepancy between the simulation and flight test response showed significant improvement, which satisfies the regulation tolerance.

Numerical analysis of high-lift configurations with oscillating flaps

This study presents two-dimensional aerodynamic investigations of various high-lift configuration settings concerning the deflection angles of droop nose, spoiler and flap in the context of enhancing the high-lift performance by dynamic flap movement. The investigations highlight the impact of a periodically oscillating trailing edge flap on lift, drag and flow separation of the high-lift configuration by numerical simulations. The computations are conducted with regard to the variation of the parameters reduced frequency and the position of the rotational axis. The numerical flow simulations are conducted on a block-structured grid using Reynolds Averaged Navier Stokes simulations employing the shear stress transport k-omega turbulence model. The feature Dynamic Mesh Motion implements the motion of the oscillating flap. Regarding low-speed wind tunnel testing for a Reynolds number of 0.5 times 10^{6} the flap movement around a dropped hinge point, which is located outside the flap, offers benefits with regard to additional lift and delayed flow separation at the flap compared to a flap movement around a hinge point, which is located at 15 % of the flap chord length. Flow separation can be suppressed beyond the maximum static flap deflection angle. By means of an oscillating flap around the dropped hinge point, it is possible to reattach a separated flow at the flap and to keep it attached further on. For a Reynolds number of 20 times 10^6, reflecting full scale flight conditions, additional lift is generated for both rotational axis positions.

A Review of Current Research in Subscale Flight Testing and Analysis of Its Main Practical Challenges

Testing of untethered subscale models, often referred to as subscale flight testing, has traditionally had a relatively minor, yet relevant use in aeronautical research and development. As recent advances in electronics, rapid prototyping and unmanned-vehicle technologies expand its capabilities and lower its cost, this experimental method is seeing growing interest across academia and the industry. However, subscale models cannot meet all similarity conditions required for simulating full-scale flight. This leads to a variety of approaches to scaling and to other alternative applications. Through a literature review and analysis of different scaling strategies, this study presents an overall picture of how subscale flight testing has been used in recent years and synthesises its main issues and practical limitations. Results show that, while the estimation of full-scale characteristics is still an interesting application within certain flight conditions, subscale models are progressively taking a broader role as low-cost technology-testing platforms with relaxed similarity constraints. Different approaches to tackle the identified practical challenges, implemented both by the authors and by other organisations, are discussed and evaluated through flight experiments.

To the Estimation of Force and Pitching Moment Coefficients Acting on the Model, Full-Scale Aircraft and Their Elements at Supercritical Angles of Attack and in a Spin

In order to clarify the results of experimental studies with free-spinning aircraft models in a vertical wind tunnel, the analysis of the force pitching moment coefficients acting on the model, full-scale aircraft and their elements was carried out. It is shown that some differences in these coefficients are acceptable, which is confirmed by a satisfactory convergence of the results of model studies and full-scale aircraft flight tests.

Design and Implementation of a Hardware-in-the-Loop Simulation System for a Tilt Trirotor UAV

The tilt trirotor unmanned aerial vehicle (UAV) is a novel aircraft that has broad application prospects in transportation. However, the development progress of the aircraft is slow due to the complicated control system and the high cost of the flight experiment. This work attempts to overcome the problem by developing a hardware-in-the-loop (HIL) simulation system based on a heavily developed and commercially available flight simulator X-Plane. First, the tilt trirotor UAV configuration and dynamic model are presented, and the parameters are obtained by conducting identification experiments. Second, taking the configuration of the aircraft into account, a control scheme composed of the mode transition strategy, hierarchical controller, and control allocation is proposed. Third, a full-scale flight model of the prototype is designed in X-Plane, and an interface program is completed for connecting the autopilot and X-Plane. Then, the HIL simulation system that consists of the autopilot, ground control station, and X-Plane is developed. Finally, the results of the HIL simulation and flight experiments are presented and compared. The results show that the HIL simulation system can be an efficient tool for verifying the performance of the proposed control scheme for the tilt trirotor UAV. The work contributes to narrowing the gap between theory and practice and provides an alternative verification method for the tilt trirotor UAV.

Research on analytical scaling method and scale effects for subscale flight test of blended wing body civil aircraft

Although a subscale flight test (SFT) is an essential step in the practical application of a blended wing body (BWB) civil aircraft, research on similarity scaling methods and the scale effects during a BWB SFT is still lacking. Therefore, this paper presents an analytical scaling method based on the standard atmospheric model and full-scale flight envelope for analysing the scale factor and flight altitude during the SFT of a BWB aircraft that uses the NPU-300 concept under different similarity conditions. The relations between the size of the subscale test vehicle (STV) and the similarity conditions are summarised, and the corresponding test scope is given. The results showed that under the premise of Froude number similarity, the minimum STV sizes were 51.54% and 75.19% of the full-scale aircraft to simulate Reynolds and Mach numbers, respectively. Extending the test altitude helped to relax the size requirements and broaden the test scope. Furthermore, a quantitative computational fluid dynamics (CFD) study of the scale effects and an analysis of the STV mass requirements were conducted for both high-speed and low-speed BWB SFTs. For the high-speed case, the scale effects mainly influenced the viscosity, surface velocity, and shock wave characteristics. The wing loading was the core parameter for the high-speed STV mass design. A 30% scaled STV with constant wing loading could control the lift–drag ratio loss to within 2.0. For the low-speed case, the scale effects changed the aerodynamic performance by reducing the boundary layer anti-separation ability and control surface efficiency. These coupling effects showed that the scale factor should be at least 50% to obtain the full-scale stall characteristics. The research on the analytical scaling method and scale effects reported in this paper can provide an approach and basis for BWB SFT design and data correction.

GPS Multireceiver Direct Position Estimation for Aerial Applications

Modern aviation safety increasingly depends on reliable GPS services, while signal degrading effects such as multipath and masking often occur during critical flight phases, such as takeoff and landing. In this regard, we propose multireceiver direct position estimation (MR-DPE), which operates a network of DPE receivers to enhance GPS accuracy under degraded signal conditions. A DPE receiver directly estimates navigation solutions in the position-velocity-time domain with a maximum-likelihood approach, bypassing the intermediate range measurements. Whereas prior works have shown the enhanced accuracy of DPE with weak signals, MR-DPE provides further improvement by leveraging the information redundancy and the geometric diversity provided by the network of receivers and antennas. We implemented MR-DPE using software-defined radio and tested it with simulated GPS signals to show improved GPS accuracy under degraded environments. We conducted comprehensive, full-scale flight experiments, a first for DPE-related works. A wide range of flight profiles was explored and analyzed, especially those prone to signal multipath and masking, thus validating the claimed benefits of MR-DPE in GPS accuracy.

Longitudinal and lateral aerodynamic characterisation of reflex wing Unmanned Aerial Vehicle from flight tests using Maximum Likelihood, Least Square and Neural Gauss Newton methods

ABSTRACTIn this paper, longitudinal and lateral-directional aerodynamic characterisation of the Cropped Delta Reflex Wing (CDRW) configuration–based unmanned aerial vehicle is carried out by means of full-scale static wind-tunnel tests followed by full-scale flight testing. A predecided set of longitudinal and lateral/directional manoeuvres is performed to acquire the respective flight data, using a dedicated onboard flight data acquisition system. The compatibility of the acquired dynamics is quantified, in terms of scale factors and biases of the measured variables, using Kinematic consistency check. Maximum likelihood (ML), least squares and newly emerging neural Gauss–Newton (NGN) methods were implemented for a wing-alone delta configuration, mainly to capture the dynamic derivatives for both longitudinal and lateral directional cases. Estimated damping and weak dynamic derivatives, which are in general challenging to capture for a wing alone configuration, are consistent using ML and NGN methods. Validation of the estimated parameters with aerodynamic model is performed by proof-of-match exercise and are presented therein.

RESEARCH TECHNIQUE OF UNMANNED AERIAL VEHICLE DYNAMIC BEHAVIOR IN GROUND CONDITIONS WITH ACCOUNT OF FLIGHT EXPERIMENT DATA

The authors of this article propose a research technique of unmanned aerial vehicle (UAV) dynamic behavior in ground conditions with account of flight experiment data. Dynamic researches of UAV and separate components of its onboard equipment are one of the most important tasks of the ground testing stage of UAV. The flight experiment gives the most reliable information about the vibration condition of the UAV and its onboard equipment. However to identify the real vibration condition of the UAV requires the installation of a sufficiently large number of accelerometers, which technically can be unrealizable during flight studies. It is proposed to solve this problem by conducting ground dynamic tests. Having determined the nature and level of vibration during the full-scale (flight) experiment in the "reference" places, it is possible further in the laboratory with a high degree of reliability to obtain vibrations (amplitude and phase frequency characteristics, spectrum, etc.) in other places of the UAV construction. To do this, it is necessary to install accelerometers throughout the UAV airframe, including the "reference" places that were used in the flight experiment. By exciting the UAV alternately along the longitudinal and lateral axes with the vibration spectrum obtained in the full-scale experiment, the vibration spectrum is reproduced in those places where accelerometers are installed in the laboratory experiment. In this case, the vibration task is carried out by "reference" accelerometers, as in the full-scale experiment. In the future, the resulting vibration process can be transferred to autonomous studies of the automatic control system and its individual units, actuator, rudder – drive system, etc. to assess their functioning in these vibration conditions. In addition to the main provisions of the technique, the article presents a schematic diagram and an example of the dynamic behavior research of UAV in ground conditions with account of flight experiment data.

Reconstruction of Atmosphere, Trajectory, and Aerodynamics for the Low-Density Supersonic Decelerator Project

The Supersonic Flight Dynamics Test was a full-scale flight test of aerodynamic decelerator technologies developed by the Low-Density Supersonic Decelerator technology development project. The purpose of the project was to develop and mature aerodynamic decelerator technologies for landing large mass payloads on the surface of Mars. The technologies included an attached toroidal Supersonic Inflatable Aerodynamic Decelerator, a ballute, and supersonic parachutes. Two flight tests were conducted at the Pacific Missile Range Facility. Sufficient data were acquired to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics for these two flight tests. This paper describes the instrumentation, analysis techniques, acquired data, and reconstruction results from these two flight tests. A technique for processing gimballed inertial measurement unit data is developed in which the inertial flight path of the instrument is reconstructed, which is then mapped back to the vehicle body frame as a postprocessing step.

Higher Harmonic Control: A Historical Perspective

Higher harmonic control (HHC) is an approach for achieving reduced helicopter vibration by controlling the vibratory rotor airloads in such a way that the fuselage excitation is minimized. This paper is a historical look at how a program aimed at helicopter vibration reduction started as an outgrowth of fixed wing flutter suppression at NASA Langley Research Center, proved the HHC concept on aeroelastically scaled wind tunnel models and went on to demonstrate viability in full-scale flight testing on the OH-6A helicopter in 1982. Following the OH-6A flight tests, the helicopter research community was stimulated to prove the effectiveness of HHC on different configurations through analysis, wind tunnel tests, and flight tests. All of these investigations have shown HHC to be effective in reducing vibration to levels not attainable with conventional vibration control methods and without any detrimental side effects. HHC development has progressed to the point that the technology provides one more option to address the ever-present vibration problem in helicopters. The literature demonstrates that helicopter ride quality equivalent to that of fixed wing aircraft is attainable with application of HHC.

Comprehensive Analysis of a Coaxial Ultralight Rotorcraft and Validation with Full-Scale Flight-Test Data

The results of full-scale flight testing in combination with comprehensive analysis of an ultralight counterrotating coaxial rotorcraft are presented. The CoAX 2D was developed and manufactured by edm aerotec in Geisleden, Germany. The helicopter features two-bladed, 6.5-m-diameter upper and lower rotors in teeter configuration with uniform airfoil sections and linearly twisted blades. The focus was on the development of a comprehensive analysis model using CAMRAD II, its validation with measured flight-test data, and results evaluation from both measurements and analysis. The CoAX 2D's power consumption, tail loads, hub loads, teeter angles, and blade-tip clearance were evaluated for both hover and forward flight. Performance and thrust correlated well between the analysis and the measurements. Effects from rotor-fuselage interference dominated the measured 2/rev hub load and 1/rev teeter angle harmonics. The numerical model also matched well with the measured vibrations during forward flight in terms of hub loads, teeter angles, and resulting blade-tip clearance. Minimal rotor-rotor blade-tip clearance occurred at low flight speeds because of dominating effects from rotor-rotor interference.

MISSION PROFILE AND DESIGN CHALLENGES FOR MARS LANDING EXPLORATION

Abstract. An orbiter and a descent module will be delivered to Mars in the Chinese first Mars exploration mission. The descent module is composed of a landing platform and a rover. The module will be released into the atmosphere by the orbiter and make a controlled landing on Martian surface. After landing, the rover will egress from the platform to start its science mission. The rover payloads mainly include the subsurface radar, terrain camera, multispectral camera, magnetometer, anemometer to achieve the scientific investigation of the terrain, soil characteristics, material composition, magnetic field, atmosphere, etc. The landing process is divided into three phases (entry phase, parachute descent phase and powered descent phase), which are full of risks. There exit lots of indefinite parameters and design constrain to affect the selection of the landing sites and phase switch (mortaring the parachute, separating the heat shield and cutting off the parachute). A number of new technologies (disk-gap-band parachute, guidance and navigation, etc.) need to be developed. Mars and Earth have gravity and atmosphere conditions that are significantly different from one another. Meaningful environmental conditions cannot be recreated terrestrially on earth. A full-scale flight validation on earth is difficult. Therefore the end-to-end simulation and some critical subsystem test must be considered instead. The challenges above and the corresponding design solutions are introduced in this paper, which can provide reference for the Mars exploration mission.

Numerical Simulations of the F-16XL at Flight-Test Conditions Using Delayed Detached-Eddy Simulation

The F-16XL flight-test data have been used for many years in the Cranked-Arrow Wing Aerodynamics Program International series as a challenging test case for numerical prediction of military aircraft at full-scale flight conditions. The complexity of the configuration and the variety of flight-test conditions available challenge most computational fluid dynamics codes and turbulence models. This study documents the use of the Kestrel flow solver with its solution-based mesh refinement capability and delayed detached-eddy simulation turbulence models for F-16XL flight condition 25, which is a subsonic high-angle-of-attack flight condition that has proven particularly challenging for computational researchers. The spectral content of the pressure data and the integrated lift, drag, and pitching moment coefficients are also presented. The more advanced turbulence models, coupled with the more refined meshes, show very good agreement with flight-test data and reasonable resolution of the unsteady frequency content.

Microphone Array Measurement in European Transonic Wind Tunnel at Flight Reynolds Numbers

This paper presents an experimental investigation of small-scale model airframe noise at various Reynolds numbers up to the real flight case. The study consists of data acquired with a microphone array in the European Transonic Wind Tunnel. The advantage of this wind tunnel is to enable testing simultaneously at cryogenic temperatures and increased pressure levels, which extends the range of achievable Reynolds numbers up to those pertaining to full-scale flight conditions. At the DLR, German Aerospace Center, the microphone array measurement technique has been further developed to perform measurements under combined cryogenic and pressurized conditions. For this purpose, a microphone array consisting of 96 microphones was designed and constructed. In this paper, aeroacoustic results are presented for various Reynolds numbers up to the real flight case on an Airbus K3DY half-model of scale . The results showed significant Reynolds number and Mach number dependency for various sources. Of particular note are various dominant sources appearing on the flap at real flight Reynolds numbers. To the author’s knowledge, this is the first time that airframe noise data for a small-scale model have been acquired at real flight Reynolds numbers.