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This paper reports an experimental study on the supersonic jet surface flow structure visualization and shear stress field measurement issuing from a rectangular nozzle with extended shelf. The evolution of the near-field surface flow structures with an increased Nozzle Pressure Ratio (NPR) is successfully captured by the surface oil flow, infrared detection technology, and the Shear-Sensitive Liquid Crystal Coating (SSLCC) technique. Results reveal that under smaller NPR, the wall flow structure is similar to that of a jet without the extended shelf i.e., clean jets, and this is caused by insufficient effect on the boundary layer. However, at higher amplitudes of NPR, there exists a significant effect of the boundary layer, as a near triangular separation forms on the trailing edge of the Mach stem due to the adverse pressure gradient, which is visualized for the very first time in this paper. Furthermore, the vector field of shear stress is measured quantitatively by SSLCC technique. Results shows that the magnitude of shear stress heightened with NPR increasing, and the directions of shear stress changes across the shock wave and expansion fans. In addition, surface streamlines measured by SSLCC is significantly consistent with the streamlines visualized using the oil flow technique.

A MXene-GaN-MXene based multiple quantum well photodetector was prepared on patterned sapphire substrate by facile drop casting. The use of MXene electrodes improves the responsivity and reduces dark current, compared with traditional Metal-Semiconductor-Metal (MSM) photodetectors using Cr/Au electrodes. Dark current of the device using MXene-GaN van der Waals junctions is reduced by three orders of magnitude and its noise spectral intensity shows distinct improvement compared with the traditional Cr/Au–GaN–Cr/Au MSM photodetector. The improved device performance is attributed to low-defect MXene-GaN van der Waals interfaces. Thanks to the high quality MXene-GaN interfaces, it is possible to verify that the patterned substrate can locally improve both light extraction and photocurrent collection. The measured responsivity and specific detectivity reach as high as 64.6 A/W and 1.93 × 1012 Jones, respectively, making it a potential candidate for underwater optical detection and communication. The simple fabrication of MXene-GaN-MXene photodetectors spearheaded the way to high performance photodetection by combining the advantages of emerging 2D MXene materials with the conventional III-V materials., A MXene-GaN van der Waals Metal-Semiconductor Junctions based High Performance Multiple Quantum Well Photodetector was demonstrated, which shows great potential for underwater wireless optical communication and turbidity sensing.

In this paper a novel design method of aerodynamic configuration is proposed to integrate forebody, strut and inlet for strutjet engine, and a model at design point of Mach number 6 is generated to investigate the aerodynamic performance by both simulations and experiments. The basic flow field employed by proposed method is a combined flow named IBB, which is combined by Internal Conical Flow A (ICFA), truncated Busemann flow I (BI) for external section, and truncated Busemann flow II (BII) for internal section. The model configuration is generated by streamline tracing method from basic flow field, in which the forebody section is traced from ICFA and BI flows, and the inlet as well as strut section is traced from BII flow. The simulations in Mach number 4, 5, and 6 demonstrate uniform starting flow fields with relatively high total pressure recovery, which agree well with experiments in wind tunnel. Additionally, in low Mach number cases, this inlet could start at Mach number 3 while it is unstarted at Mach number 2.7; in high Mach number cases, a uniform flow could still exist in Mach number 6.5 while a relatively strong shock wave boundary layer interaction is found in cowl area of Mach number 7 case, indicating the inlet designed by proposed method works in a relatively wide Mach number range.

The internal drag of the unconventional inner flow channel of blended wing body aircraft must be measured accurately to correct the air intake effect of the blended wing body flow-through model in wind tunnel tests. In this study, the pressure distribution of the inner flow channel under the interaction of internal and external flows was obtained through numerical simulation. A specialized pressure-measuring rake was designed based on the numerical results, and a validation test was conducted in a 2.4 m × 2.4 m transonic wind tunnel. Compared with the flow in traditional inlets/nozzles, the flow in the unconventional inner channel in the current research is asymmetric, the distortion index is higher, and the internal drag is more sensitive to flow changes. The wind tunnel test results have a good correlation with the numerical results, and the repeatability of the test results is satisfactory, indicating that the measurement accuracy and precision of the pressure-measuring rake are acceptable. The design method of the specialized rake is feasible, and it can be used to guide the measurement of complex flow in unconventional inner flow channels of blended wing body aircraft.

A better understanding of the mechanism of flow separation suppression by a dielectric barrier discharge is essential for flow control. This paper investigates the mechanism of improving the aerodynamic performance of the airfoil by dielectric barrier discharge when the Reynolds number is in the range of 6 × 104–4 × 105. The results show that the disturbance of the gas discharge to the flow field will form a new flow structure. The fluctuating frequency of the new flow structure determines the ability of the plasma actuator to suppress flow separation. This investigation improves and develops the mechanism of plasma flow control.

In this paper, the incident shock–separation shock interactions on a surface plate near the leading edge are studied theoretically and experimentally, and the transition from regular reflection (RR) to Mach reflection (MR) is the main focus. The theoretical method employs free interaction theory (FIT) and the minimum entropy production (MEP) principle to analyse the separation shock strength of flow separated from the boundary layer and separated from the leading edge, respectively, the criterion based on the MEP principle is employed to predict the RR-to-MR transition near the leading edge. The experiments were performed on a rotatable wedge situated over a sharp leading-edge plate such that the wedge could continuously change the flow deflection angle from by means of a high-precision control device. Fast-response transducers and a high-speed camera were used to measure dynamic pressures and to take schlieren images, respectively. The influences of wedge positions, Reynolds numbers and Mach numbers on shock reflections are investigated by careful tests. The theoretical and experimental results for Mach numbers 5, 6, 7 and 8 show good agreement, indicating that the theoretical method is applicable.

In this work, the Direct Numerical Simulation (DNS) and Oil-Film Interferometry (OFI) technique are used to investigate the hypersonic boundary layer transition induced by single and double roughness elements at Mach number 5. For single roughness, the DNS results showed that both horseshoe vortices and hairpin vortices caused by shear layer instability can affect the boundary layer instability. The generation of the near-wall unstable structure is the key point of boundary layer transition behind the roughness element. At the downstream of the roughness element, the interaction between horseshoe vortices and hairpin vortices will spread in the spanwise direction. For double roughness elements, the effect of the spacing between roughness elements on the transition is studied. It is found that the case of higher spacing between roughness elements is more effective for inducing transition than the lower one. The interaction between two adjacent roughness elements can suppress the evolution of horseshoe vortices in the downstream and trigger the instability of shear layer. Thus, the transition will be suppressed accordingly. Keywords: Boundary layer transition, Hypersonic, Direct numerical simulation (DNS), Oil-film interferometry, Roughness elements

The parallel modular inlet is a common configuration in a strutjet engine of the rocket-based combined-cycle hypersonic vehicle. To study the flow interaction between two adjacent inlets, the flowf...

Prediction of boundary layer transition is important for the design of hypersonic aircrafts. The study of boundary layer transition of hypersonic flow around a flat plate using oil-film interferometry was investigated at Φ500mm traditional hypersonic wind tunnel. In order to measure the skin friction fast and precisely on the hypersonic wind tunnel, the traditional oil-film interferometry technique is improved. A high-speed camera is used to capture the images of fringes and the viscosity of the silicon oil is modified according to the wall temperature measured by thermocouples during the test. The skin frictions of smooth surface and the surface with single square roughness element were measured. For the smooth surface, the boundary layer is laminar. However, the boundary layer transition is promoted by wake vortices induced by the roughness element. Both the results of skin friction with and without the roughness element are in good agreement with the simulation results correspondingly, indicating high accuracy of the oil film interferometry technique.

In this paper the configurations of shock wave-boundary layer interactions (SWBLI) are studied theoretically and experimentally in Mach number 2 and 2.5 flows on test models with various wedge angles ranging from to. The proposed theoretical method couples the free interaction theory (FIT) with the minimum entropy production (MEP) principle to predict the appearance of separation shock, resulting in convex, straight and concave separation shock waves according to different solution combinations, which agree well with current experiments. Additionally, several influences on SWBLI are studied experimentally, in which the parameters related to theoretical solutions are found mostly determining the flow configuration, and SWBLI is much more sensitive to incident shock strength than incoming flow properties. Separation could be suppressed by incident shock when the MEP solution is smaller than the FIT, while it could be intensified when the MEP solution is larger than FIT; by contrast, the effects of separation position and model mounting height could be very weak.

In this paper separation-induced shock reflection is studied theoretically and experimentally. An analytical model is proposed to establish the connections among upstream conditions, downstream conditions and shock configurations. Furthermore, the minimum entropy production principle is employed to determine the incident shock angles as well as the criterion for the transition from regular reflection to Mach reflection, which agrees well with experimental results. Additionally, a solution path for a reflected shock that fulfills the minimum entropy production principle is found in the overall regular reflection domain, based on which the steadiest shock configuration may be determined according to upstream and downstream conditions.

The aim of present investigation is to analyze the unsteady oblique shock train and boundary layer interactions during the self-excited and forced oscillation. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to the forced oscillation. A Schlieren system as well as transient pressure measurements and particle image velocimetry have been used to capture quantitative and qualitative shock structure information. Results show that the behaviors of unsteady SBLIs structure are highly related to the dynamics of shock motion. For both self-excited oscillation and forced oscillation, the asymmetrical characteristics of first X-shock was found to be negatively correlated with shock velocity. There exist some relative motions between the first X-shock and the second shock, but the absolute variations are very weak. At lower excitation frequency, the relative motion is not noticeable to the oscillation amplitude, it could be treated as a rigid motion in the duct. At higher excitation frequency, the relative motion amplitude is significant to the oscillation amplitude, and the relative movement of shock cells becomes the dominant motion. There is a hysteretic effect and phase lag between the shock position and downstream pressure perturbation when the shock train travels along a different path for upstream and downstream movements, and the hysteretic effect becomes weaker with increasing frequency.

An experimental investigation of the response of an oblique shock train to downstream periodic pressure perturbations was conducted. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to a periodic oscillatory motion of the oblique shock train. Six cases of perturbation frequency are studied. The results indicate that the downstream pressure perturbations propagate upstream to cause the oblique shock train to oscillate with a translational motion back and forth and the wall pressure to fluctuate with the same frequency. There is no distinct relative motion between the first oblique shock and the second shock during the motion process of the oblique shock train. The entire oblique shock train exhibits its behaviour of rigid motion and the strength of the first oblique shock of the oblique shock train is nearly stable during its periodic motion. There is a clear hysteresis effect in that the oblique shock train travels along a different path for the upstream and downstream motions. A simple analytical model was built based on these experimental data to analyse the oblique shock train dynamics.

The surface dielectric barrier discharge plasma actuator driven by nanosecond pulses is recognized as an effective fluid actuator for flow separation control. The operation condition of nanosecond dielectric barrier discharge actuators for separated flow control still requires further study, particularly prioritizing the improvement of the effectiveness and reducing energy consumption in flow separation control implementation. In this study, experiments are conducted using a two-dimensional NASA SC(2)-0712 airfoil in a wind tunnel with a Reynolds number of 0.5 × 106 (25 m/s). The pressure measurement experiments show that the location of actuators affects the efficiency of separation control. Particle image velocimetry results indicate that the most efficient location of the actuator is upstream of the separation point and near the original point of the separated shear layer. Meanwhile, the particle image velocimetry results show the vorticity attaches to the airfoil wall after discharge, which suggests that the reattachment is due to the generation of large-scale vortices. These present structures result in the mixing of the shear layer with the main flow thereby delaying separation and reattaching a separated flow. This study shows the most efficient location related to the separation point. Furthermore, it indicates the reattachment of flow is attributed to the motion of vortexes coherent structure.

The modern high performance air vehicles are required to have extreme maneuverability, which includes the ability of controlled maneuvers at high angle of attack. However, the nonlinear and unsteady aerodynamic phenomena, such as flow separation, vortices interaction, and vortices breaking down, will occur during the flight at high angle of attack, which could induce the uncommanded motions for the air vehicles. For the high maneuverable and agile air missile, the nonlinear roll motions would occur at the high angle of attack. The present work is focused on the self-induced nonlinear roll motion for a missile configuration and discusses the influence of the strake wings on the roll motion according to the results from free-to-roll test and PIV measurement using the models assembled with different strake wings at α = 60°. The free-to-roll results show that the model with whole strake wings (baseline), the model assembled with three strake wings (Case A) and the model assembled with two opposite strake wings (Case C) experience the spinning, while the model assembled with two adjacent strake wings (Case B), the model assembled with one strake wing (Case D) and the model with no strake wing (Case E) trim or slightly vibrate at a certain “×” rolling angle, which mean that the rolling stability can be improved by dismantling certain strake wings. The flow field results from PIV measurement show that the leeward asymmetric vortices are induced by the windward strake wings. The vortices would interact the strake wings and induce crossflow on the downstream fins to degrade the rolling stability of the model. This could be the main reason for the self-induced roll motion of the model at α = 60°.

This paper is aimed at the problem of the aircraft’s post stall maneuver, simulate aircraft precursor with tangent-ogive cylinder and utilize pitch oscillation mechanism to simulate the dynamic process of the pitching motion of aircraft. Static and dynamic experiments were investigated in a low speed wind tunnel in NUAA. Quantitative PIV test results, model’s surface pressure distribution and leeward flow field were obtained. Experimental study shows: both symmetric and asymmetric flow field exist behind the slender body in the range of high angle of attack. When the angle of attack is 30°, the initial position of the asymmetric flow field is at 3.5D section of the slender body along the axis. When the angle of attack increases to 35°, the initial position of the asymmetric flow field moves to the 2.5D section position. The changing of the structure, intensity and spatial position of the leeward vortex during the dynamic pitching motion leads to the lateral and normal force hysteresis. Relative to the static state, the asymmetrical flow field position changed in the axial direction of the slender body during the dynamic pitching motion. The initial angle of attack of generating lateral force at the 3.5D cross section changed. In the upward pitching motion, the asymmetric flow field moved to the downstream and the initial angle of attack delayed from 30° to 40°. In the downward pitching, the asymmetric flow field moved upstream and the initial angle of attack is advanced from 30° to 20°.

A cruciform-finned slender body can excite a limit-cycle rolling oscillation at high angles of attack. To suppress the unwanted motions, flow control approaches should be employed if the aerodynamic control surfaces lose control efficiency at high angles of attack. As a promising technology, the ns-dielectric barrier discharge plasma actuator has been successfully used in high-speed and high-Reynolds-number flow control applications. The present work employs a π-type ns-dielectric barrier discharge plasma actuator and vortex generators to suppress self-excited rolling oscillation at α = 50°. The free-to-roll tests show that the plasma actuator ignited at F+ ∼ 1.5 and that the vortex generators can suppress rolling oscillation. The flow patterns from particle image velocimetry measurement at different cross-sections and rolling angles suggest that the vorticity decrease of the leeward vortices may be the control mechanism for the plasma actuator. For the vortex generators, evident modification of the flow field structure can be observed due to the vortices generated from the vortex generators, which decreases the rolling moment induced by the asymmetry vortices to suppress the self-excited rolling oscillation.

Next-generation aircraft and missile are required to have extreme maneuverability, which should maintain stability at high angle of attack. However, the unsteady flow field surrounding the air vehicle would affect the aerodynamics loads and induce unwanted nonlinear motions, of which the rolling motions are usually generated. In order to investigate the unsteady rolling characteristics of a cruciform-finned slender body, the free-to-roll and force measurement tests including particle image velocimetry measurement have been conducted comprehensively. Different types of rolling motion, trimming at equilibrium positions, self-excited rolling oscillation, and self-excited spinning are observed during the free-to-roll experiment depending on the different angles of attack. The force measurement results show that the rolling motions are related to both the static and dynamic stabilities of the rolling moment at balance points. The stabilities of the rolling moment would change as the angle of attack for the model increases. At last, the flow field results from particle image velocimetry measurement indicate that the unsteady rolling motions may be induced by the interaction between the asymmetry vortices and strake wings, fins.

The basic dynamic characteristic of a nanosecond (NS) pulse discharge actuator in still air is investigated, which introduces viscid and inviscid effects to the flow, namely, heated air and pressure waves, respectively. The efficacy of a dielectric barrier discharge plasma actuator driven by a nanosecond pulse for active flow separation control is investigated experimentally on a NASA SC(2)-0712 aerofoil at Re = 0.5 × 10 6 (25 m/s). The pressure distribution on the aerofoil surface is measured, and the result indicates that the separation is well controlled at a high angle of attack. The PIV test result shows that the forcing frequency selectively affects the flow over the aerofoil, and a higher forcing frequency has a better influence on the leading edge shear layer. Schlieren imaging experiments at Re = 0.1 × 10 6 (5 m/s) show that the main disturbance factor for the nanosecond pulse plasma actuator is the heating effect, which changes the local temperature and density and induces several vortex structures. The vortex structures entrain the outer high-momentum fluid into the shear layer, thereby strengthening the mixing of the shear layer with the main flow and delaying separation or even reattaching a separated flow. The research reveals the major disturbance factor introduced by NS discharge applied to flow separation control. Furthermore, the spatiotemporal disturbance process induced by the pulse discharge is investigated.

A new dynamic test mechanism has been developed in the unsteady low-speed wind tunnel at Nanjing University of Aeronautics and Astronautics to enhance the abilities of experiments in the unsteady flow cases. The structure of the new test mechanism is described in detail in this paper and the equations used to control the performance of test model’s motions are also listed here. Then a wing-body model was tested in the new test system for the first time, including the forced oscillation tests and free-to-roll test at high angle of attack region. The test results are also discussed.

An experimental study is reported of supersonic jet surface flow structure visualization and wall shear stress field measurement issuing from a rectangular nozzle with aft-deck. The near-field surface flow structures evolution from over-expansion to under-expansion with the increase of nozzle pressure ratio (NPR) are successfully captured by surface oil flow visualization and shear sensitive liquid crystal coating (SSLCC) technique. The quantitative measurement result of shear stress vector field obtained by SSLCC shows that shear stress directions change significantly across the shock wave and expansion fans, while the magnitudes of shear stress have no obvious changes. Surface streamlines calculated by SSLCC image keep great consistency with the streamlines visualized using oil flow technique, which demonstrates the accuracy and potential application of SSLCC in supersonic jet surface flow visualization.

In this study, we investigated the effects of surface sliding discharge of a dielectric barrier discharge (SL-DBD) plasma actuator on the boundary layer flow using the time-resolved particle image velocimetry (PIV) technique. According to the PIV measurements, the peak-shaped mean jet flow is induced by the SL-DBD plasma actuator in quiescent air. Under the freestream of U∞ = 6 m/s on the flat plate, the results show that the streamwise velocity fluctuation induced by SL-DBD plasma is increased pronouncedly at the downstream of the actuator. Meanwhile, the mean flow velocity increases obviously in the downstream of the actuator. The instantaneous results show the formation of the vorticity at the bottom of the boundary layer after the plasma excitation. In addition, the vorticity has the behavior to lift up, possibly because of the induced peak-shaped jet.

An optical fiber displacement sensor based on the microwave photonics interferometric (MWPI) method is proposed and experimented, which provides an ideal solution for large range displacement measurement with high resolution. The sensor used a Michelson microwave photonics interferometer to sense the displacement with one sensing arm and a length-adjusted reference arm. The displacement variation would change the period of the microwave response function of the interferometer. According to the principle that the phase difference in one free spectral range (FSR) of the microwave response function is 360&deg, the displacement can be retrieved by the microwave response function by means of a vector network analyzer (VNA). A programmable path-switching true time delay line was used in the reference arm to decrease the microwave bandwidth. The measurement results show that the displacement sensing range is larger than 3 m and the measurement resolution is 31 &mu, m. Finally, the measurement stability is tested, and the factors affecting the measurement resolution of this method and the main source of errors are investigated in detail.

The oblique shock train (OST) flow diffusion phenomenon, which involves an interaction between the duct’s peripheral boundary layer and central oblique-shock-wave field, usually appears in constant or nearly constant cross-sectional-area supersonic/hypersonic duct flows. The study of such a complex flow structure in a confined duct under a finite adverse pressure gradient has important implications for the design and operation of a variety of devices including hypersonic vehicle inlet/isolator [1, 2], supersonic wind tunnel diffusers [3], supersonic ejectors, and so on. To develop the design methods and control strategies of the flow devices, it is necessary to fully understand the mechanism of the OST. Different flow conditions will lead to different kinds of shock train structure: normal shock train occurs for lower incoming Mach numbers of about 1.2–2, whereas oblique shock train occurs for higher incoming Mach numbers. The aerodynamic performance of the fluid devices is well related to the features of the OST, which include the structure of the OST, the pressure recovery in the OST flow region, the dynamic characteristics of the flow separation region nearby the OST, and so on. In addition to complex flow structure, the OST inside the duct has an important feature of unsteady characteristics; even if the boundary conditions such as the pressures upstream and downstream of the OST region are held constant, the OST is still in “self-oscillation” state [4]. This self-excited oscillation may cause the structural fatigue, unstart of inlet/isolator, and instability of combustion. Therefore, the study of dynamic characteristics of the OST is important for aircraft structural design and aerodynamic performance evaluation.

In order to improve quality and efficiency of Blended Wing-Body (BWB) aircraft’s high-speed wind tunnel test, while verify the reliability and practicability of flange balance in the test of such a layout aircraft. The specific flange-connected balance of BWB aircraft was designed and processed, which was designed under the limits of aerodynamic load and model interspace size. The comparison test results between the flange balance and the traditional cone balance were obtained by conducting a comparison test in the 2.4-meters transonic wind tunnel. The results show that the flange balance has high measurement accuracy and good correlation with the traditional cone balance. Flange balance meets the requirement of BWB layout aircraft high speed wind tunnel test and has a broad application prospect, so that it could be popularized in the future.

Detailed phase-lock schlieren measurement was conducted to study the physical evolution process of the structures inside the plasma synthetic jet actuator driven by a high voltage nanosecond pulse. One strong blast wave and heat source core are generated immediately after the arcing discharge. Within the cavity, as the blast wave is reflected by the cavity walls and interacts with the heat source core, the pressure of the gas inside the cavity is increased rapidly and a series of diffractive wave comes out from the orifice of the cavity, followed by a train of vortex rings and a high speed jet.

According to the latest drilling data, the geologic survey of surrounding outcrops and the laboratory test results of rock samples, this paper studies the regional distribution, geochemical and reservoir characteristics and gas concentration of organic-rich shales to reveal the shale gas generation and potential of the Lower Cambrian Qiongzhusi Formation of the Southern Sichuan Basin. The study area features high organic content (with TOC between 0.55% and 25.7%, average of more than 2%), great net shale thickness (of 60–300 m), high brittle mineral content (over 40%), abundant micro- and nanometer-sized pores and micro-fractures, and high gas content (0.27–6.02 m3/t, average of 1.90 m3/t), which is in favor of the generation and accumulation of shale gas. The Lower Cambrian Qiongzhusi Formation is one of the most favorable strata which has frequent gas show and industrialization breakthrough during drilling, and the Weiyuan and Xunyong-Junlian areas are two most realistic and favorable targets for shale gas exploration and development. Key words: Southern Sichuan Basin, Lower Cambrian, Qiongzhusi Formation, shale gas, accumulation, resource potential

With Sichuan Basin as focus, this paper introduces the depositional environment, geochemical and reservoir characteristics, gas concentration and prospective resource potential of three different types of shale in China: marine shale, marine-terrigenous shale and terrigenous shale. Marine shale features high organic abundance (TOC: 1.0%–5.5%), high-over maturity (Ro: 2%–5%), rich accumulation of shale gas (gas concentration: 1.17–6.02 m3/t) and mainly continental shelf deposition, mainly distributed in the Paleozoic in the Yangtze area, Southern China, the Paleozoic in Northern China Platform and the Cambrian-Ordovician in Tarim Basin; Marine-terrigenous coalbed carbonaceous shale has high organic abundance (TOC: 2.6%–5.4%) and medium maturity (Ro: 1.1%–2.5%); terrigenous shale in the Mesozoic and Cenozoic has high organic abundance (TOC: 0.5%–22.0%) and mid-low maturity (Ro: 0.6–1.5%). The study on shale reservoirs in the Lower Paleozoic in Sichuan Basin discoveried nanometer-sized pores for the first time, and Cambrian and Silurian marine shale developed lots of micro- and nanometer-sized pores (100–200 nm), which is quite similar to the conditions in North America. Through comprehensive evaluation, it is thought that several shale gas intervals in Sichuan Basin are the practical targets for shale gas exploration and development, and that the Weiyuan-Changning area in the Mid-South of Sichuan Basin, which is characterized by high thermal evolution degree (Ro: 2.0%–4.0%), high porosity (3.0%–4.8%), high gas concentration (2.82–3.28 m3/t), high brittle mineral content (40%–80%) and proper burial depth (1500–4500 m), is the core area for shale gas exploration and development, the daily gas production for Well Wei 201 is 1×104–2×104 m3. 摘要: ： 以四川盆地为重点，介绍中国海相、海陆过渡相、陆相三大类型页岩形成的沉积环境、地球化学与储集层特征、含气量与远景资源量。中国海相页岩是一套高有机质丰度（TOC为1.0%～5.5%）、高—过成熟（Ro值为2.0%～5.0%）、富含页岩气（含气量1.17～6.02 m3/t）、以陆棚相为主的沉积，主要分布在华南扬子地区古生界、华北地台古生界和塔里木盆地寒武系—奥陶系；海陆过渡相煤系炭质页岩有机质丰度高（TOC为2.6%～5.4%）、成熟度适中（Ro值为1.1%～2.5%）；中新生界陆相页岩有机质丰度高（TOC为0.5%～22.0%）、低熟—成熟（Ro值为0.6%～1.5%）。在对四川盆地下古生界页岩储集层研究中首次发现，寒武系和志留系海相页岩发育大量与北美地区相似的微米—纳米级孔隙。综合评价认为四川盆地发育的多套页岩气层系是勘探开发的现实领域，四川盆地中南部威远—长宁等地区的寒武系和志留系是页岩气勘探开发的核心区与层系，其特点是：热演化程度较高（Ro值为2.0%～4.0%）、孔隙度较高（3.0%～4.8%），含气量较高（2.82～3.28 m3/t）、脆性矿物含量较高（40%～80%）、埋深适中（1 500～4 500 m），有利于开采。图7表7参38 Key words: unconventional hydrocarbon, shale gas, nanometer-sized pore throat, shale oil, tight oil, source rock hydrocarbon, core area

The influence of distributed cylinder roughness elements on boundary layer transition for NACA0012 airfoil at Ma = 0.6 has been investigated by subsonic/transonic/supersonic wind tunnel experiment with oil-film interferometry. Three different heights and two different distances of cylinder roughness elements on the airfoil model were used, and the skin friction coefficient was measured by the oil-film interferometry. The experimental results show that higher roughness elements promote the transition earlier. In addition, narrower distance of roughness elements can delay the transition compared with the case of wider distance.

In order to obtain the flow field characteristics and the influence of boundary layer, numerical simulations and wind tunnel tests are conducted for two streamline traced Jaws inlets at Mach number 7. The inlets are designed based on a flow field with 8-7 planar shock wave (the ramp in pitch plane is inclined at 8° to the free stream and in yaw plane is inclined at 7° to the free stream, yielding planar shocks). In the study, the static pressure distributions were measured and analyzed along the plane-symmetric centerline of the inlet with and without the boundary layer correction, respectively. Results show that boundary layer correction can obviously weaken the viscous influence to the inlet, increasing the mass flow coefficient and improving total pressure recovery.

The dynamics of three vortices moving in an ideal fluid in a plane can be expressed in Hamiltonian form. However, when viscosity can not be ignored, the system of point vortices has not this delicious structure. This investigation focuses on the viscosity effect on the motion of three vortices. Since the viscous diffusion of vortices is simulated by Brownian motion, the equations of intervortical distances are obtained by Itǒo formula. These equations show that there exist two viscosity effects on intervortical distances: stable expansion and random impulse. Furthermore, via employing the asymptotic solution to these equations, these two viscous effects have been shown to destroy the self-similar process of collapse and make it into a new configuration, which is similar to the near collapse in the ideal case.