Naoko Tokugawa

Transition measurements on the natural laminar flow wing at Mach 2

Transition characteristics of a supersonic natural laminar flow wing with a sweep angle of more than 60 deg were measured using hot-film sensors and an infrared camera at Mach 2. The pressure falls very rapidly in the vicinity of the leading edge of the wing, such that crossflow instabilities are reduced. The flow slightly accelerates downstream such that streamwise instabilities are stabilized. Tests were conducted in two wind tunnels to estimate the transition location in flight: a quiet tunnel with a low Reynolds number and a moderately quiet tunnel with variable Reynolds number. As a result, the natural laminar flow effect of the wing was confirmed. The unit Reynolds number had no effect on the transition Reynolds number of the wing. The transition Reynolds number was 8 x 10 5 at the design angle of attack at Mach 2, but there is room for further investigation

Transition Measurement of Natural Laminar Flow Wing on Supersonic Experimental Airplane NEXST-1

layer. In this paper, the results of the transition measurement are introduced and compared with the numerically predicted results. The transition locations detected experimentally are in good agreement with the predicted locations, and the natural laminar flow effect is confirmed in the aerodynamic design conditions of the supersonic experimental airplane NEXST-1. Nomenclature C = local chord length CL = lift coefficient of full configuration Cp = surface pressure coefficient Cprms = fluctuation (rms) of surface static pressure of the wind tunnel E = dc output of the hot-film sensor e = ac output of the hot-film sensor H = altitude M = Mach number p = local surface pressure measured by the dynamic pressure transducer PPRT = local total pressure measured by the Preston tube Rec = Reynolds number based on the mean aerodynamic chord S = semispan length Tblow = time from the beginning of the blow in the windtunnel test Tlo = time from the liftoff in the flight test TTC = local temperature measured by the thermocouple X = chordwise position Xtip = axial length from the tip of the wind-tunnel model Y = spanwise position � = angle of attack ’ = circumferential angle from the top line on the nose 0 = fluctuation

Transition Detection on Supersonic Natural Laminar Flow Wing in the Flight

The experimental validation of natural laminar flow wing designed with the original CFDbased inverse design method is carried out by the flight test of an unmanned and scaled supersonic experimental airplane. It is the first challenge to apply the natural laminar flow wing concept to a supersonic vehicle. The concept of natural laminar flow wing is validated by measuring the surface pressure and the transition location. In this paper, the transition characteristics detected by the hot-film sensors and dynamic pressure transducers are discussed. Transition location is classified by using newly introduced quantity, called “transition level”, based on objective criteria. The transition location detected experimentally is in good agreement with numerically predicted location, and the natural laminar flow effect is confirmed at the design condition.

Experiments on Streamline-Curvature Instability in Boundary Layers on a Yawed Cylinder

Instability of the three-dimensional boundary layer on a yawed circular cylinder placed in a uniform flow is investigated experimentally by introducing acoustic disturbances from a point near the attachment line. To exemplify the flow dominated by streamline-curvature instability rather than crossflow instability, which has been often observed in many swept-wing flows, is the aim here. In upstream regions of the disturbance wedge originating from the point source, both streamline-curvature and crossflow disturbances are superposed on each other and yield complicated amplitude distributions. A newly proposed method enables the decomposition of the distorted amplitude distribution into contributions from the two instability modes. Detailed observations, however, show that the crossflow mode decays with the distance from the source much faster than the streamline-curvature mode and allows the latter to be dominant in a region further downstream. A fundamental characteristic of the streamline-curvature instability wave is confirmed by examining its phase distribution in the spanwise and normal directions.

Pressure Gradient Effects on Supersonic Transition over Axisymmetric Bodies at Incidence

Boundary-layer transition on axisymmetric bodies at a nonzero angle of attack in Mach 2 supersonic flow was investigated using experimental measurements and linear stability analysis. Transition over four axisymmetric bodies (namely, the Sears–Haack body, the semi-Sears–Haack body, the straight cone, and the flared cone) with different axial pressure gradients was measured in two different facilities with different unit Reynolds numbers. The semi-Sears–Haack body and flared cone were designed specifically to achieve a broader range of axial pressure distributions. Measurements revealed a dramatic effect of body shape on transition behavior near the leeward plane of symmetry. For a body shape with an adverse pressure gradient (that is, a flared cone), the measured transition patterns show an earlier transition location along the leeward symmetry plane in comparison with the neighboring azimuthal locations. For a nearly zero pressure gradient (that is, the straight cone), such leeward-first transition is ob...

Boundary-layer transition on an axisymmetric body at incidence at Mach 1.2

Experimental investigation of the boundary-layer transition on an axisymmetric nose model at 1- and 2-deg incidence was conducted at Mach 1.2. The configuration of the model is the forward part of a Sears-Haack body defined to have minimum wave drag caused by volume at 0-deg incidence in supersonic flow. Transition locations were obtained with small surface roughness on the order of 0.1 μ using an infrared camera. An unsteady pressure transducer was applied to investigate instability mechanisms that lead to transition. These results show that the most aft transition occurred on the leeward ray rather than on the windward rays as observed on sharp cones. Tollmien-Schlichting instability dominated the transition process on the windward ray, and crossflow instability was assumed to dominate on the side similar to the transition on the sharp cones. However, transition occurred more aft on the leeward ray than for the case of the sharp cones, and the transition front was determined by turbulent wedges formed as a consequence of the more forward transition on the side. The disturbance believed to be a traveling crossflow wave was observed at 1-deg incidence, and its measured frequency was in good agreement with the maximum amplified frequency, which was calculated using a compressible linear e N code.

Transition Within Leeward Plane of Axisymmetric Bodies at Incidence in Supersonic Flow

Boundary layer transition along the leeward symmetry plane of axisymmetric bodies at nonzero angle of attack in supersonic flow was investigated experimentally and numerically as part of joint research between the Japan Aerospace Exploration Agency (JAXA) and National Aeronautics and Space Administration (NASA). Transition over four axisymmetric bodies (namely, Sears-Haack body, semi-Sears-Haack body, straight cone and flared cone) with different axial pressure gradients was measured in two different facilities with different unit Reynolds numbers. The semi-Sears-Haack body and flared cone were designed at JAXA to broaden the range of axial pressure distributions. For a body shape with an adverse pressure gradient (i.e., flared cone), the experimentally measured transition patterns show an earlier transition location along the leeward symmetry plane in comparison with the neighboring azimuthal locations. For nearly zero pressure gradient (i.e., straight cone), this feature is only observed at the larger unit Reynolds number. Later transition along the leeward plane was observed for the remaining two body shapes with a favorable pressure gradient. The observed transition patterns are only partially consistent with the numerical predictions based on linear stability analysis. Additional measurements are used in conjunction with the stability computations to explore the phenomenon of leeward line transition and the underlying transition mechanism in further detail.

Experimental investigation of the flow instability near the attachment-line boundary layer on a yawed cylinder

The behavior of small disturbances in a 3-D laminar boundary layer on a yawed cylinder was experimentally investigated. This setup simulates the flow around the leading edge of swept wings. Since multiple instability modes appear near the attachment-line region, a point-source disturbance was artificially introduced to separate these modes. Amplitude and phase distributions of the disturbances originating from the point source were measured using a hotwire probe near the attachment-line flow to test existing theoretical predictions. Hotwire measurements show that two instability modes definitely coexist and overlap in the middle portion of the wedge-shaped region developing downstream of the point source. Decomposition by 2-D fast Fourier transform (FFT) analysis enables us to separate one oblique wave from the other. One of the oblique waves belongs to the cross-flow instability mode, which travels to the attachment line and grows even at Reynolds numbers that are slightly lower than the critical Reynolds number for the attachment-line instability. The origin of the other mode is not identifiable, because it has peculiar characteristics different from both the streamline-curvature instability mode and the cross-flow instability mode. This mode decays in the downstream direction for all frequencies examined. By investigating the spatial characteristics of the small disturbance, the critical Reynolds number for cross-flow instability was successfully determined in the off-attachment-line region. The value, Rc = 543, was lower than the critical Reynolds number of Rc = 583 for the attachment-line flow. Furthermore, the critical frequency and wavenumber were in good agreement with existing predictions based on linear stability theory.

Excitation of streamline-curvature instability in three-dimensional boundary layer on a yawed cylinder

Instability of the three-dimensional boundary layer on a yawed circular cylinder placed in a uniform-flow is investigated experimentally, by introducing acoustic disturbances from a point near the attachment line. The aim is to exemplify the flow cbminated by streamlinecurvature instability, rather than cross-flow instability which is often observed in many swept-wing flows. In upstream regions of the disturbance wedge originating from the source point, both streamlinecurvature and cross-flow disturbances arc superposed on each other to yield so complicated results of hot-wire measurements that cannot be compared with theoretical pre&tions. Detailed observations, however, show that the crossflow mcde decaysmuch faster with the distance from the source and allows the streamline-curvature mode to bc dominant in a far &wnstmam region. Some fundamental characteristicsof the two instability waves are confirmed by examining the phase distributions in the spanwise and normal directions. from the usual process governed by C-F streamwise vortices. On the other hand, we have very poor knowledge of S-C instability in boundary layers on swept wings, becausethe phenomenon has beenignored or unrecognizeduntil the recentthcoretical prediction by Itoh and the following experimental confirmation by Takagi et al.” It is highly desirable to investigate. roles of SC instability in the transitional process of such wing flows. As a first step of such investigations, the present study is drected to the pmblem of whether SC disturbance can have a magnitude comparable to or larger than C-F mo& or not. Fortunately, linear stability theory has predrcted some geometrical conditions under which S-C m& is unstable but C-F mode is always stable in the swept-cylinder flow’. Artificial excitation is made to clarify the characteristics of the flow by acoustically intro&ing a point-source disturbance through a tiny hole near the leading edgeof the model.

Characteristics of streamline-curvature disturbances in a rotating-disk flow

An experimental investigation of rotating-disk flow is made to reveal the fundamental characteristics of a centrifugal instability caused by curvature of inviscid external streamlines. This was recently shown to be identical to the so-called Type n or parallel instability hi atmospheric sciences. Small but clearly observable disturbances of this instability are successfully produced by means of artificial excitation through the annular slit of the disk in a low Reynolds number region, where crossflow instability is subcriticaL Acoustic forcing simulating various environmental fluctuations is also used to investigate contribution of the streamlinecurvature instability to transition. Experimental results show that streamline-curvature disturbances initiated at the slit behave like annular waves travelling in the outboard direction, whose characteristics are in good agreement with linear stability theory. Anotherimportant result is that the introduction of a weak streamlinecurvature wavepacketoverwhelmspre-existing cross-flow stationary vortices, suggesting that a gust or turbulent patch hi surrounding fluid may cause the transition location to shift downward. On the contrary, continuously strong forcing of streamline-curvature instability leadsto another transition route, which is differentfrom the usual transition dominated by streamwise vortices.