The influence of non-equilibrium dissociation on the flow produced by shock impingement on a blunt body

We describe an investigation of the effects of non-equilibrium thermochemistry on the interaction between a weak oblique shock and the strong bow shock formed by a blunt body in hypersonic flow. This type of shock-on-shock interaction, also known as an Edney type IV interaction, causes locally intense enhancement of the surface heat transfer rate. A supersonic jet is formed by the nonlinear interaction that occurs between the two shock waves and elevated heat transfer rates and surface pressures are produced by the impingement of the supersonic jet on the body. The current paper is motivated by previous studies suggesting that real gas effects would significantly increase the severity of the phenomenon. Experiments are described in which a free-piston shock tunnel is used to produce shock interaction flows with significant gas dissociation. Surprisingly, the data that are obtained show no significant stagnation enthalpy dependence of the ratio of the peak heat transfer rates with and without shock interaction, in contrast to existing belief. The geometry investigated is the nominally two-dimensional flow about a cylinder with coplanar impinging shock wave. Holographic interferometry is used to visualize the flow field and to quantify increases in the stagnation density caused by shock interaction. Time-resolved heat transfer measurements are obtained from surface junction thermocouples about the model forebody. An improved model is developed to elucidate the finite-rate thermochemical processes occurring in the interaction region. It is shown that severe heat transfer intensification is a result of a jet shock structure that minimizes the entropy rise of the supersonic jet fluid whereas strong thermochemical effects are promoted by conditions that maximize the entropy rise (and hence temperature). This dichotomy underlies the smaller than anticipated influence of real gas effects on the heat transfer intensification. The model accurately predicts the measured heat transfer rates. Improved understanding of the influence of real gas effects on the shock interaction phenomenon reduces a significant element of risk in the design of hypersonic vehicles. The peak heat transfer rate for the Edney type IV interaction is shown to be well-correlated, in the weak impinging shock regime, by an expression of the form [equation] for use in practical design calculations

Engineering Fluid Dynamics 2019-2020

This book contains the successful submissions to a Special Issue of Energies entitled “Engineering Fluid Dynamics 2019–2020”. The topic of engineering fluid dynamics includes both experimental and computational studies. Of special interest were submissions from the fields of mechanical, chemical, marine, safety, and energy engineering. We welcomed original research articles and review articles. After one-and-a-half years, 59 papers were submitted and 31 were accepted for publication. The average processing time was about 41 days. The authors had the following geographical distribution: China (15); Korea (7); Japan (3); Norway (2); Sweden (2); Vietnam (2); Australia (1); Denmark (1); Germany (1); Mexico (1); Poland (1); Saudi Arabia (1); USA (1); Serbia (1). Papers covered a wide range of topics including analysis of free-surface waves, bridge girders, gear boxes, hills, radiation heat transfer, spillways, turbulent flames, pipe flow, open channels, jets, combustion chambers, welding, sprinkler, slug flow, turbines, thermoelectric power generation, airfoils, bed formation, fires in tunnels, shell-and-tube heat exchangers, and pumps

Application of integrated fluid-thermal-structural analysis methods

Hypersonic vehicles operate in a hostile aerothermal environment which has a significant impact on their aerothermostructural performance. Significant coupling occurs between the aerodynamic flow field, structural heat transfer, and structural response creating a multidisciplinary interaction. Interfacing state-of-the-art disciplinary analysis methods is not efficient, hence interdisciplinary analysis methods integrated into a single aerothermostructural analyzer are needed. The NASA Langley Research Center is developing such methods in an analyzer called LIFTS (Langley Integrated Fluid-Thermal-Structural) analyzer. The evolution and status of LIFTS is reviewed and illustrated through applications

Underexpanded Supersonic Plume Surface Interactions: Applications for Spacecraft Landings on Planetary Bodies

Numerical and experimental investigations of both far-field and near-field supersonic steady jet interactions with a flat surface at various atmospheric pressures are presented in this paper. These studies were done in assessing the landing hazards of both the NASA Mars Science Laboratory and Phoenix Mars spacecrafts. Temporal and spatial ground pressure measurements in conjunction with numerical solutions at altitudes of approx.35 nozzle exit diameters and jet expansion ratios (e) between 0.02 and 100 are used. Data from steady nitrogen jets are compared to both pulsed jets and rocket exhaust plumes at Mach approx.5. Due to engine cycling, overpressures and the plate shock dynamics are different between pulsed and steady supersonic impinging jets. In contrast to highly over-expanded (e 5 (lunar atmospheric regime), the ground pressure is minimal due to the development of a highly expansive shock structure. We show this is dependent on the stability of the plate shock, the length of the supersonic core and plume decay due to shear layer instability which are all a function of the jet expansion ratio. Asymmetry and large gradients in the spatial ground pressure profile and large transient overpressures are predominantly linked to the dynamics of the plate shock. More importantly, this study shows that thruster plumes exhausting into martian environments possess the largest surface pressure loads and can occur at high spacecraft altitudes in contrast to the jet interactions at terrestrial and lunar atmospheres. Theoretical and analytical results also show that subscale supersonic cold gas jets adequately simulate the flow field and loads due to rocket plume impingement provided important scaling parameters are in agreement. These studies indicate the critical importance of testing and modeling plume-surface interactions for descent and ascent of spacecraft and launch vehicles

Experimental Investigations for Heat Transfer Characteristics of Under-expanded Impinging Jet

Department of Mechanical EngineeringIn this study, under-expanded impinging gas jets are investigated experimentally for understanding heat transfer characteristics of the jets. As working fluids, nitrogen (N2) and carbon dioxide (CO2) are selected in order to observe heat transfer effect changed by different working fluid. CO2 jet has a potential way to enhance the heat transfer effect which is sublimation. The novel concept of dry-ice assisted jet impingement cooling is proposed in this study. When carbon-dioxide (CO2) passes through a tiny orifice gap or jet nozzle, it experiences a rapid temperature drop as well as a pressure decrease via the Joule-Thomson effect. Joule-Thomson coefficient of CO2 is proven to be higher than the coefficient of other gases??? such as nitrogen, hydrogen, air. This temperature drop causes the formation of small CO2 dry-ice particles. In addition to the enhanced cooling performance caused by lowered bulk-jet temperature, heat transfer is improved by the additional sublimation effect between the dry-ice particles and the cooling target surface. A comparison of the cooling performance between the suggested CO2 solid-gas two-phase jet and a single-phase nitrogen (N2) jet was performed experimentally as well. In order to form dry-ice particles, high pressure and velocity of jet fluid are inevitably required, which is enough for compressible effect to appear. Both jets have differences not only in phase change but also in jet flow structures. As jet velocity increased, shock structures at jet downstream appeared and surface temperature is changed as well. The structures are detected more clearly in N2 jet than CO2 jet because of the difference in total pressure at jet boundary, therefore, relationship between shock structure of N2 jet flow and heat transfer is investigated in this study. In case of high Reynolds number impinging jet which is enough to reach supersonic flow regime, stagnation temperature of impinged surface is affected by jet structure as well as other factors such as nozzle-plate distance or radial distance does. When the jet flow velocity becomes supersonic, shock structures are constructed at downstream of the nozzle exit. Complicated shock structure such as Mach shock disk, plate shock is highly expected to affect to the heat transfer behavior of impingement surface. In this study, a cooling performance of supersonic N2 jet is investigated by measuring the impinged surface temperature and the flow of the jet is visualized by Schlieren image system. Visualized image and surface temperature are compared to clarify the flow structure-related heat transfer characteristics. In all the experiments of present study, jet fluids are expanded through a circular nozzle and impinged on an electrically heated flat heater surface, and their heat transfer coefficients are measured. The performances of the impinging jet for both fluids are also evaluated via the variance of flow parameters, for example, the Reynolds number, and the jet geometry configurations.ope

Heat removal in high pressure turbine seal segments

An important parameter for turbomachinery designers is “clearance control”, because the clearances between interfaces must be set to optimum values to maximize power output, operational life and efficiency. Leakage of hot gas result- ing from excessive clearance, can lead to flow instabilities, components overheat- ing, lower cycle efficiency and a dramatic increase in specific fuel consumption (SFC). Seal segments are used to reduce blade tip leakage, maintain coolant air flow and the stability of rotor-dynamic systems, helping to maximize blade perfor- mance. Seal segments in the High-Pressure Turbine (HPT) stages are one of the hottest components as they face the hot gases coming from the combustion chamber with temperatures which can reach 1700 0 C and which makes them sub- ject to oxidation, erosion, and creep. Thus, seal segments need to be protected. They are currently cooled using jet impingement techniques, passing cooling air (supplied by the high-pressure stage of the compressor) through channels to di- rectly impinge on the hot surfaces. The focus of this research was to improve the jet impingement cooling of the seal segments in HPTs by investigating methods that provide more effective heat removal. The role played by configurations of ribs (surface roughness using be- spoke turbulators), custom-made seal-segments, and surface features such as contouring, both in isolation and combination, were investigated using numerical methods. A set of 174 simulations were carried including the use of uniform and non-uniform roughness elements with different shapes and heights. Firstly, three different uniform roughness elements were tested, a square cross-sectional continuous rib, a hemi-spherical pin-fin and a cubical pin-fin for three jet impingement angles of α=90°, 60° and 45°. Each roughness element was also tested for six different heights (e) between 0.25 mm and 1.5 mm in increments of 0.25 mm. Results are presented in the form of average Nusselt number within and beyond the stagnation region. Secondly, the effect of using a roughness element with a square cross section in the shape of a circle, on the average Nu was investigated for four different radial locations (R), three jet angles (α) and six rib heights (e). Finally, the roughness element used was continuous, of square cross-sec- tion, in the shape of tear drops and reversed tear drops. This meant the rib did not act as a total barrier to flow in either the uphill or downhill direction

Experimental studies on shock wave interactions with flexible surfaces and development of flow diagnostic tools

Nowadays, light-weight composite materials have increasingly used for high-speed flight vehicles to improve their performance and efficiency. At supersonic speed, sonic fatigue, panel flutter, severe instabilities, and even catastrophic structural failure would occur due to the shock wave impingement on several flexible components of a given structural system either internally or externally. Therefore, investigation on shock wave interaction with flexible surfaces is crucial for the safety and performance of high-speed flight vehicles. This work aims to investigate the mechanism of shock wave interaction with flexible surfaces with and without the presence of the boundary layer. The first part involves the shock wave generated by supersonic starting jets interaction with flexible surfaces and the other one focuses on shock wave and boundary layer interaction (SBLI) over flexible surfaces. A novel miniature and cost-effective shock tube driven by detonation transmission tubing was designed and manufactured to simulate the supersonic starting jet and investigate the interaction of a supersonic starting jet with flexible surfaces. To investigate the characterization of this novel type shock tube, the pressure-time measurement in the driven section and the time-resolved shadowgraph were performed. The result shows that the flow structure from the open end of the shock tube driven by detonation transmission tubing agrees with that of conventional compressed-gas driven shock tubes. Moreover, this novel type of shock tube has good repeatability of less than 3% with a Mach number range of 1.29-1.58 when the weight of the NONEL explosive mixture varies from 3.6mg to 12.6mg. An unsteady background oriented schlieren (BOS) measurement system and a sprayable Polymer-Ceramic unsteady pressure sensitive paint (PC-PSP) system were developed. The preliminary BOS result in a supersonic wind tunnel shows that the sensitivity of the BOS system is good enough to visualize weak density variations caused by expansion waves, boundary layer, and weak oblique shocks. Additionally, compared with the commercial PC-PSP from Innovative Scientific Solutions Incorporated (ISSI), the in-house developed unsteady PSP system has higher pressure sensitivity, lower temperature sensitivity, and photo-degradation rate. To identify the shock movement, distortion and unsteadiness during the processes of the supersonic starting jet impingement and shock wave boundary layer interaction (SBLI) over flexible surfaces, an image processing scheme involving background subtraction in the frequency domain, filtering, resampling, edge detection, adaptive threshold, contour detection, feature extraction, and fitting was proposed and applied to process shadowgraph and schlieren sequences automatically. A large shadowgraph data set characterized by low signal to noise ratio (SNR) and small spatial resolution (312×260-pixel), was used to validate the proposed scheme. The result proves that the aforementioned image processing scheme can detect, track, localize, and fit shock waves in a subpixel accuracy. The mechanism of the interaction between the initial shock wave from a supersonic starting jet and flexible surfaces was investigated based on a square shock tube driven by detonation transmitting tube. Compared with that of the solid plate case, flexible surfaces can delay the shock reflection process because of the flexible panel deformation generated by the pressure difference between the top and the bottom. The delay time is around 8µs in the case of 0.1mm thick flexible surface, whereas it declines to around 4µs in the case of 0.3mm thick flexible surface because of the lower flexibility and deformation magnitude. However, interestingly, the propagation velocity of the reflected shock wave is basically the same for the solid plate and flexible panels, which means the flexible surface doesn’t reduce the strength of the reflection wave, although it delays its propagation. Also, there is not an apparent difference in the velocity of the reflected shock wave in the case of different incident shock Mach numbers when Ms varying from 1.22 to 1.54. These experimental results from this study are useful for validating numerical codes that are used for understanding fluid-structure interaction processes

Shock Interaction Control for Scramjet Cowl Leading Edges

An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface

Publications in acoustics and noise control from the NASA Langley Research Center during 1940 - 1974

This document contains reference lists of published Langley Research Center papers in various areas of acoustics and noise control for the period 1940-1974. The research work was performed either in-house by the center staff or by other personnel supported entirely or in part by grants or contracts. The references are listed chronologically and are grouped under the following general headings: (1) Duct acoustics, (2) Propagation and operations, (3) Rotating blade noise, (4) Jet noise, (5) Sonic boom, (6) Flow-surface interaction noise, (7) Human response, and (8) Structural response