Free-flight Measurements Research Articles

Wind tunnel measurements of dynamic aerodynamic coefficients using a freely rotating test bench

PurposeThis paper aims to propose a dedicated measurement methodology able to simultaneously determine the stability derivative Cmα̇ and the pitch damping coefficient sum Cmq + Cmα̇ in a wind tunnel using a single and almost non-intrusive metrological setup called MiRo.Design/methodology/approachTo assess the MiRo method’s reliability, repeatability and accuracy, the measurements obtained with this technique are compared to other sources like aerodynamic balance measurements, alternative wind tunnel measurements, Ludwieg tube measurements, free-flight measurements and computational fluid dynamics (CFD) simulations. Two different numerical approaches are compared and used to validate the MiRo method. The first numerical method forces the projectile to describe a pure oscillation motion with small amplitude along the pitch axis during a rectilinear flight, whereas the second numerical approach couples the one degrees of freedom simulation motion equations with CFD methods.FindingsMiRo, a novel and almost non-intrusive technique for dynamic wind tunnel measurements, has been validated by comparison with five other experimental and numerical methodologies. Despite two completely different approaches, both numerical methods give almost identical results and show that the holding system has nearly no impact on the dynamic aerodynamic coefficients. Therefore, it could be assessed that the attitude of MiRo model in the wind tunnel is very close to the free-flight one.Originality/valueThe MiRo method allows studying the attitude of a projectile in a wind tunnel with the least possible impact on the flow around a model.

Free-Flight Aerodynamic Testing of the Skylon Space Plane

The determination of aerodynamic coefficients for complex high-speed vehicles still requires experimental measurement. This paper details the development of the free-flight measurement technique within the University of Oxford High Density Tunnel. In particular, a novel image processing technique is developed for calculating the position of the model from high-speed video. This study focuses on the measurement of high Reynolds number experimental aerodynamic data for a subscale model of the Skylon space plane of Reaction Engines. Testing was undertaken at a Mach 7 test condition replicating flight at an altitude of 63.5 km. Results for lift, drag, and pitching moment coefficients were obtained over a range of angles of attack. Lift coefficient was nearly linear over the range of angles of attack tested, and drag coefficient was parabolic in shape, but sensitive to model yaw. The vehicle was also shown to be statically unstable, a common characteristic of canard configuration vehicles.

Aerodynamic characteristics of a free-flight scramjet vehicle in shock tunnel

A multi-component aerodynamic test for an airframe-engine integrated scramjet vehicle model was conducted in the free-piston shock tunnel HIEST. A free-flight force measurement technique was applied to the scramjet vehicle model named MoDKI. A new method using multiple piezoelectric accelerometers was developed based on overdetermined system analysis. Its unique features are the following: (1) The accelerometer’s mounting location can be more flexible. (2) The measurement precision is predicted to be improved by increasing the number of accelerometers. (3) The angular acceleration can be obtained with single-axis translational accelerometers instead of gyroscopes. (4) Through the averaging process of the multiple accelerometers, model natural vibration is expected to be mitigated. With eight model-onboard single-axis accelerometers, the three-component aerodynamic coefficients (Drag, Lift, and Pitching moment) of MoDKI were successfully measured at the angle of attack from 0.7 to 3.4 degrees under a Mach 8 free-stream test flow condition. A linear regression fitting revealed a 95% prediction interval as the measurement precision of each aerodynamic coefficient.Graphical abstract

Aerodynamic characterization of space debris in the VKI Longshot hypersonic tunnel using a free-flight measurement technique

Hypersonic experiments are performed in the VKI Longshot wind tunnel on hemispherical and annular geometries representative of space debris, aiming at determining their lift, drag and pitching moment coefficients. A free-flight technique is used for this purpose where the model is automatically released in the hypersonic flow and safely recovered by the end of the experiment. A schlieren flow visualization technique coupled with a high-speed camera follows the motion of the object as it is exposed to the high-velocity flow. The instantaneous attitude of the free-flying model and its trajectory are determined from image analysis techniques. Correction terms are included to account for the gravity and for the relative motion of the object within the flow. Experimental results are compared with both numerics and theory.

Accurate position control of a flapping-wing robot enabling free-flight flow visualisation in a wind tunnel

Flow visualisations are essential to better understand the unsteady aerodynamics of flapping wing flight. The issues inherent to animal experiments, such as poor controllability and unnatural flapping when tethered, can be avoided by using robotic flyers that promise for a more systematic and repeatable methodology. Here, we present a new flapping-wing micro air vehicle (FWMAV)-specific control approach that, by employing an external motion tracking system, achieved autonomous wind tunnel flight with a maximum root-mean-square position error of 28 mm at low speeds (0.8–1.2 m/s) and 75 mm at high speeds (2–2.4 m/s). This allowed the first free-flight flow visualisation experiments to be conducted with an FWMAV. Time-resolved stereoscopic particle image velocimetry was used to reconstruct the three-dimensional flow patterns of the FWMAV wake. A good qualitative match was found in comparison to a tethered configuration at similar conditions, suggesting that the obtained free-flight measurements are reliable and meaningful.

Sensitivity Analysis for Global Parameter Identification. Application to Aerodynamic Coefficients

The parameter identification of grey-box nonlinear model structures remains a challenging task. Analytical models, constructed from physical laws, allow a complete description of the system behavior, with however a high complexity of the mathematical structure of the system. This paper presents an identification procedure adapted to nonlinear models and able to cope with the difficulties related to the model structure. This procedure, composed of several steps from model development to its validation, takes into account the identifiability of model parameters. It highlights the importance of performing sensitivity analysis of the output variables w.r.t. unknown parameters - input conditions and/or model parameters - to give insights on the choice of the experimental conditions able to provide sufficient information for the estimation step. These results are illustrated in the case of the identification of aerodynamic coefficients of a space probe based on free flight measurements.

Aerodynamic Coefficient Identification of a Space Vehicle from Multiple Free-Flight Tests

The identification of aerodynamic coefficients, based on free-flight measurements, remains complex and challenging for vehicles such as space probes, unmanned aerial vehicles, or ammunition. In this paper, a detailed procedure for the identification of the drag, pitching moment slope, and pitch damping coefficients of a reentry space vehicle is presented. New models of these aerodynamic coefficients, based on polynomial functions of Mach numbers and incidence angles, are proposed, and an estimation strategy using multiple data series is applied. The free-flight data correspond to three-axis magnetometer and radar measurements for different experimental conditions. The resulting model is tested with new data sets to ascertain the model validity. The results show that the proposed description of the three aerodynamic coefficients and the estimation strategy are relevant.

Free-flight measurement technique in the free-piston high-enthalpy shock tunnel.

A novel multi-component force-measurement technique has been developed and implemented at the impulse facility JAXA-HIEST, in which the test model is completely unrestrained during the test and thus experiences free-flight conditions for a period on the order of milliseconds. Advantages over conventional free-flight techniques include the complete absence of aerodynamic interference from a model support system and less variation in model position and attitude during the test itself. A miniature on-board data recorder, which was a key technology for this technique, was also developed in order to acquire and store the measured data. The technique was demonstrated in a HIEST wind-tunnel test campaign in which three-component aerodynamic force measurement was performed on a blunted cone of length 316 mm, total mass 19.75 kg, and moment of inertia 0.152 kgm(2). During the test campaign, axial force, normal forces, and pitching moment coefficients were obtained at angles of attack from 14° to 32° under two conditions: H0 = 4 MJ/kg, P0 = 14 MPa; and H0 = 16 MJ/kg, P0 = 16 MPa. For the first, low-enthalpy condition, the test flow was considered a perfect gas; measurements were thus directly compared with those obtained in a conventional blow-down wind tunnel (JAXA-HWT2) to evaluate the accuracy of the technique. The second test condition was a high-enthalpy condition in which 85% of the oxygen molecules were expected to be dissociated; high-temperature real-gas effects were therefore evaluated by comparison with results obtained in perfect-gas conditions. The precision of the present measurements was evaluated through an uncertainty analysis, which showed the aerodynamic coefficients in the HIEST low enthalpy test agreeing well with those of JAXA-HWT2. The pitching-moment coefficient, however, showed significant differences between low- and high-enthalpy tests. These differences are thought to result from high-temperature real-gas effects.

Free Flight and Wind Tunnel Measurements of the Drag Exerted on an Archery Arrow

We describe two support-interference-free measurements of aerodynamic forces exerted on an archery arrow. The first measurement is performed in a wind tunnel with JAXA's 60cm Magnetic Suspension and Balance System, whereas the second measurement is carried out by a free flight experiment. In the latter measurement, the trajectory and rotation of the arrow is recorded by two high-speed video cameras and the arrow's down range velocity decay rate is analyzed to determine the drag coefficient. The arrow-shooting system developed by Miyazaki et al. (2013) is utilized, in order to investigate Re number dependence of the drag coefficient for . We attach two points (piles) of different type (streamlined and bullet) to the arrow-head and two kinds of vanes (SPIN- WING-VANE and GAS PRO) to the arrow-tail. The boundary layer remains laminar for any combination of point and vane, if Re is less than . It becomes turbulent for Re larger than and the drag coefficient increases to about 2.6, when the bullet point and SPIN-WING-VANE are attached. In the same Re range, two values of drag coefficient are found for the combination of streamlined point and SPIN-WING-VANE. The lower value (about 1.6) corresponds to a laminar boundary layer and the larger value (about 2.6) to a turbulent boundary layer. In contrast, for GAS PRO vanes, the boundary layer remains laminar at any Re considered, irrespective of the point- shape. These findings confirm that both the point- and vane-shapes have a crucial influence on the laminar to turbulent transition of the boundary layer. We also asked an elite female archer (recurve bow) to shoot the same A/C/E arrow (with bullet point and SPIN-WING-VANE), of which velocity is 58 m/s (). The arrow launched by the archer oscillates along its length (Archer's Paradox), and the drag coefficient is found to be about2.65 (turbulent value), indicating that the oscillation of the arrow induces the boundary layer transition.

Pressure, temperature and free-flight projectile-displacement measurements during a plasma discharge generated on a supersonic projectile

The generation of a plasma discharge on a projectile surface is a way of producing a pressure imbalance in order to divert a projectile from its initial trajectory. Thus, some experiments dealing with the steering of a supersonic projectile by using a plasma actuator are presented. The plasma discharge is produced by an embedded low-voltage generator capable of delivering a power controlled electric discharge between the electrodes flush with the projectile surface. A series of experiments carried out under wind-tunnel conditions at a Mach number of 3 allows surface-pressure and temperature measurements, which are synchronised with the visualisation of the perturbation evolution along the projectile surface, the disturbance being generated by the plasma discharge. Other series of experiments realised in a shock tunnel under 2.5 km of altitude real conditions at a Mach number of 4.5 demonstrates that a plasma discharge generates the angular deviation of an explosively-formed-projectile model: 120 J of stored energy delivered in 4.5 ms in front of the ogive-cylinder junction of the projectile induces an angle of attack of 11° after the plasma-discharge generation.

High-throughput measurement of free-flight kinematics in the dragonfly

High-throughput measurement of free-flight kinematics in the dragonfly

A wing-assisted running robot and implications for avian flight evolution

DASH+Wings is a small hexapedal winged robot that uses flapping wings to increase its locomotion capabilities. To examine the effects of flapping wings, multiple experimental controls for the same locomotor platform are provided by wing removal, by the use of inertially similar lateral spars, and by passive rather than actively flapping wings. We used accelerometers and high-speed cameras to measure the performance of this hybrid robot in both horizontal running and while ascending inclines. To examine consequences of wing flapping for aerial performance, we measured lift and drag forces on the robot at constant airspeeds and body orientations in a wind tunnel; we also determined equilibrium glide performance in free flight. The addition of flapping wings increased the maximum horizontal running speed from 0.68 to 1.29 m s−1, and also increased the maximum incline angle of ascent from 5.6° to 16.9°. Free flight measurements show a decrease of 10.3° in equilibrium glide slope between the flapping and gliding robot. In air, flapping improved the mean lift:drag ratio of the robot compared to gliding at all measured body orientations and airspeeds. Low-amplitude wing flapping thus provides advantages in both cursorial and aerial locomotion. We note that current support for the diverse theories of avian flight origins derive from limited fossil evidence, the adult behavior of extant flying birds, and developmental stages of already volant taxa. By contrast, addition of wings to a cursorial robot allows direct evaluation of the consequences of wing flapping for locomotor performance in both running and flying.

Evaporative and excretory water loss during free flight in pigeons

Body water conservation is important in flying birds because the very high metabolic demands and heat dissipation requirements during flight depend on plasma-volume integrity. Wind tunnel experiments and theoretical model predictions show that evaporative water loss (EWL) depends on air temperature (Ta) and water vapor density (ρa), but these relationships have not been examined in free-flying birds. The contribution of excretory water loss to the total water loss of a flying bird is thought to be negligible but this assumption is untested. To study the dependence of water losses on environmental conditions in free-flying birds and to quantify the contribution of excretory water loss to total water loss, we estimated evaporative and excretory water losses in 16 trained, free-flying tippler pigeons (Columba livia, 250–340 g). We collected excreta by attaching a light latex, water-impermeable receptacle around each bird's vent. By gravimetry, we measured evaporative and excretory water losses of birds for eight flights at different Tas and compared flying to resting (control) birds for two of these flights. EWL was constant with respect to Ta when less than 15 °C, and increased with increasing Ta above 19 °C, indicating that evaporative cooling was invoked when the heat load increased. EWL increased with increasing ρa, possibly due to the strong correlation between ρa and Ta. Excretory water loss was independent of ρa or Ta and averaged almost 10% of the total water loss. Measurements of EWL made on pigeons during wind tunnel experiments and previous free-flight studies are consistent with our free-flight measurements made at similar Tas.

Laser Doppler Anemometer for In-Flight Velocity Measurements on Airplane Wings

The development of laminar wing technology for commercial airplanes requires strategies that combine wind-tunnel investigations with numerical flow computations and in-flight measurements. One strategy is outlined; it is explained that instruments are needed to measure local flow velocities in wind tunnels as well as to carry out free flight experiments. It is shown that laser Doppler anemometry (LDA) is most suited to provide the required local velocity information, but special LDA systems need to be developed. Results of LDA system developments are summarized. This program yielded two optical units suitable for in-flight velocity measurements on airplane wings. Laboratory and free flight measurements were successfully carried out with the LDA systems developed, and a summary of results is presented. Applications to laminar wing design are indicated. Suggestions for further advancements of the LDA systems are presented

Density measurements in ballistic re-entry

The interdisciplinary application of experimental and numerical methods is the crucial strategy for the profound validation of numerical simulations of complex aerothermodynamic flows. Especially for the design of reusable heat shield materials the reliable re-entry heat flux prediction from numerical simulation methods became important. In this context free flight measurements gained an outstanding position for the validation of theoretical models describing the high-temperature phenomena of thermo-chemically reacting hypersonic flows in interaction with the flight vehicle's surface. This article is devoted to present a global concept for the application of a compact free flight sensor exposed to the extreme thermal environmental condition of a controlled, ballistic re-entry. The validation strategy will be demonstrated for total density measurements proposed for the stagnation point of a re-entry capsule, since the density at the capsule surface reflects sensitively the interaction process of the hot gas and the wall. The partial validation of the aerothermodynamic simulation tool was elaborated for laboratory experiments in the gun-tunnel and shock tube. The measuring principle of the optical density sensor, the Phase-compensated Michelson Interferometer (PMI), will be discussed in detail. Since the feasibility of these measurements are highly dependent on the thermal integration of the sensor, an effective thermal protection concept, based on modern ceramic matrix composite materials and high performance insulations, has been verified in numerical analyses.

AEROTHERMODYNAMIC ANALYSIS FOR AXISYMMETRIC PROJECTILES AT SUPERSONIC/HYPERSONIC SPEEDS

An aerothermodynamic design code for axisymmetric projectiles has been developed using a viscous‐inviscid interaction scheme. Separate solution procedures for the inviscid and the viscous (boundary layer) fluid dynamic equations are coupled by an iterative solution procedure. Non‐equilibrium, equilibrium and perfect gas boundary layer equations are included. The non‐equilibrium gas boundary layer equations assume a binary mixture (two species; atoms and molecules) of chemically reacting perfect gases. Conservation equations for each species include finite reaction rates applicable to high temperature air. The equilibrium gas boundary layer equations assume infinite rate reactions, while the perfect gas equations assume no chemical reactions. Projectile near‐wall and surface flow profiles (velocity, pressure, density, temperature and heat transfer) representing converged solutions to both the inviscid and viscous equations can be obtained in less than two minutes on minicomputers. A technique for computing local reverse flow regions is included. Computations for yawed projectiles are accomplished using a coordinate system transformation technique that is valid for small angle‐of‐attack. Computed surface pressure, heat transfer rates and aerodynamic forces and moments for 1.25 &le Mach No. &le 10.5 are compared to wind tunnel and free flight measurements on flat plate, blunt‐cone, and projectile geometries such as a cone‐cylinder‐flare.

Autorotating flat-plate wings: the effect of the moment of inertia, geometry and Reynolds number

Free-flight and wind-tunnel measurements by previous investigators of the flat-plate autorotation phenomenon have been analysed. The variation of the autorotation characteristics with changes in the Reynolds number and the aspect ratio, thickness ratio and moment of inertia of the flat plate have been correlated. The interpretation of the role of the Reynolds number made in a previous investigation is shown to be incorrect. The tip-speed ratio, for the ranges of the dimensionless parameters investigated, is shown to be a function of only the plate aspect ratio, thickness ratio, and also the moment of inertia if the latter is sufficiently small. The lift and drag coefficients, and therefore the free-flight glide angle, are shown to be functions of the tip-speed ratio, the aspect ratio and the Reynolds numbers based on the chord and plate thickness.

A Computer-Controlled Video Instrument for Wind-Tunnel Testing of Lifting Parachutes

Wind-tunnel tests of parachutes require that aerodynamic data be obtained simultaneously on both the forebody and the parachute. In particular, it is necessary to correlate the relative motion between the parachute and forebody to the loads exerted by the parachute on the forebody as a function of time during dynamic tests, disreefing, and for the full-open canopy. A computerized video instrumentation technique was developed to provide this correlation during a full-scale wind-tunnel test of a unique lifting parachute configuraton. The instrumentation consisted of a small, rugged TV camera, mounted on the forebody base, which tracked the positions of two lights attached to the inside of the parachute canopy. The positions of the lights were digitized by an on-line minicomputer and converted to yaw, pitch, and roll angles relative to the forebody. These data were obtained 30 times per second and stored on disk memory along with instantaneous values of the axial force and rolling moment exerted by the parachute on the forebody. The motion-force correlations made it possible to determine the yaw damping of the lifting parachute and provided information on the specific sources of rolling moments generated by the parachute. Test results indicate that this computer-controlled video instrumentation can be applied to free-flight tests, dynamic tests, and measurements of relative motion among multiple bodies.

Reply by Author to M. J. Walsh

Henderson, C. B., Coefficients of Spheres in Continuum and Rarefied Flows, AIAA Journal, Vol. 14, June 1966, pp. 707708. Bailey, A. B. and Hiatt, J., Free-Flight Measurements of Sphere Drag at Subsonic, Transonic, Supersonic, and Hypersonic Speeds for Continuum, Transition, and Near-Free-Molecular Flow Conditions, Arnold Engineering Development Center, Tullahoma, Tenn., AEDCTR-70-291, March 1971. Walsh, M. J., Coefficient Equations for Small Particles in High Speed Flows, AIAA Journal, Vol. 13, Nov. 1975, pp. 15261528. Walsh, M. J., Influence of Particle Drag Coefficient on Particle Motion in High-Speed Flow with Typical Laser Velocimetry Applications, NASATN D-8120, Feb. 1976. Zarin, N. A., Measurement of Non-Continuum and Turbulence Effects on Subsonic Sphere Drag, NASA CR-1585, June 1970. Emmons, H. W., Ed., Fundamentals of Gas Dynamics, Vol. HI. High Speed Aerodynamics and Jet Propulsion, Princeton University Press, Princeton, N.J., 1958, p. 689.

Free flight measurements of catastrophic water drop breakup

Free flight measurements of catastrophic water drop breakup