Hypersonics

In  March  2019,  we  published  an  analysis  entitled  Hypersonic  Missiles  –  a  Political  Multipurpose Weapon. It was aimed at evaluating political and strategic dimensions of the development of hypersonic weapons against the background of technical, operational and economic realities.  The basic conclusions we reached at that time are valid now as well. However, some important recent developments in this field are worth noticing. Thus, in the present analysis we will depict how the maturation of the weaponry mentioned two and a half years ago has progressed and what is new within the realm of hypersonics. We will also offer some slightly updated strategic assessments.

1.1 Hypersonic missiles are a new type of missile which possesses speed and manoeuvring capabilities that make them a game-changer in nuclear (and conventional) interstate competition1. They are much faster than existing cruise missiles and much more manoeuvrable than existing ballistic missile re-entry vehicles (RVs). They can reach and maintain hypersonic speeds, i.e. speeds exceeding Mach 52. However, this characteristic is not unique to them: ICBM RVs also travel at hypersonic speeds3. Hypersonic missiles follow a non-ballistic trajectory, flying between 18 miles and 60 miles in altitude. Their manoeuvrability allows them to change course up to the last minutes of flight4 and achieve a high degree of targeting precision. These characteristics have implications for detection. The unusual altitude and flight path of hypersonic missiles can result in their being invisible to existing missile early-warning radars for much of their trajectory. Some varieties5 will be will be detected by satellites during their boost-phase. HCMs are a two-stage weapon system, the first-stage is a small rocket booster that will accelerate the missile at the right supersonic speed needed for properly starting the second stage. During the firing of the first-stage there could be the possibility for a boost-phase detection. It is still not certain if the infrared signature of the first stage of HCM could be detected by infrared satellites used for missile warning. See JAMES ACTON (n.2), p. 68.

Hypersonic weapons have in recent years attracted attention from militaries, governments and, increasingly, multilateral institutions following reports of successful prototype testing. In 2018, the UN Secretary-General highlighted hypersonic glide vehicles and cruise missiles in a report on the role of science and technology in the context of international security and disarmament, and called on the international community to ‘remain vigilant in understanding new and emerging weapon technologies that could imperil the security of future generations’. The development of hypersonic weapons is said to pose a challenge to strategic missile defences and raise wider international security concerns due to their ‘considerable potential to further complicate strategic relations, encourage new arms competition and endanger stability’. The development and potential future deployment of hypersonic weapons illuminates a number of broader themes and questions that deserve attention from the perspective of multilateral weapons control.

This note presents the results of a shock tunnel study in which skin friction, heat transfer rates and static pressure are measured in hypervelocity turbulent boundary layers. The experiments are conducted at Mach numbers ranging from 4.2 to 5.9, flow stagnation enthalpies of 4.8 MJ/kg to 9.5 MJ/kg, and static pressures of 14 kPa to 86 kPa. The objective of the present investigation is to explore the Reynolds analogy at high stagnation enthalpy and high Mach number over a broad range of cf values. Two previous shock tunnel investigations1,2 indicate that there is a decreasing measured Reynolds analogy factor with increased measured skin friction coefficient at Mach numbers ranging from 4.6 to 7.1. The present note presents results for Reynolds analogy factors at higher skin friction coefficients than the previous studies, and also includes results with nitrogen as a test gas to suppress real-gas effects.

The HyShot 2 flight in 2002 pioneered a low cost method of hypersonic testing using sounding rockets. In a further development of this technology, Flight 7 of the HIFiRE Program will deploy a free-flying scramjet powered payload that is planned to enter the atmosphere at a high flight path angle at approximately Mach 8. The HIFiRE 7 payload consists of two back-to-back three-dimensional scramjet engines set within a trailing flare for aerodynamic stability. The scientific goals of the flight are to perform a direct measurement of the thrust generated by the scramjet flowpaths, and to compare the measured thrust with estimates based on ground testing. This paper describes the motivation, launch procedure, flowpath development and ground testing for the flight which is scheduled for March 2011.

Hypersonic technology is expected to revolutionise the way countries throughout the world can counter ballistic and nuclear attacks. With these technologies arises a new kind of arms race, putting today's global balance of power at risk-hence the importance of highlighting the pros and the cons of these increasingly coveted weapons, and the global impact they may have. Hypersonic weapons: a disruptive technology?

The object of this study is to canvas the literature for the purpose of identifying and compiling a list of Gaps, Obstacles, and Technological Challenges in Hypersonic Applications (GOTCHA). The significance of GOTCHA related deficiencies is discussed along with potential solutions, promising approaches, and feasible remedies that may be considered by engineers in pursuit of next generation hypersonic vehicle designs and optimizations. Based on the synthesis of several modern surveys and public reports, a cohesive list is formed, consisting of widely accepted areas needing improvement and falling under several general categories. These include: aerodynamics, propulsion, materials, analytical modeling, CFD modeling, and education in high speed flow physics. New methods and lines of research inquiries are suggested such as the homotopy-based analysis (HAM) for the treatment of strong nonlinearities, the use of improved turbulence models and unstructured grids in numerical simulations, the need for accessible validation data, and the refinement of mission objectives for Hypersonic Air-Breathing Propulsion (HABP).

In this thesis, an extensive study and analyze is performed to understand turbine based combined cycle engines as a propulsion system for hypersonic aircrafts. Thermal cycle was discussed and analyzed to fit the needs of HYCAT-1A a civil airplane capable of carrying 200 passengers for 9260 km, while cruising at altitude 30 km with speed six times the speed of sound.
Study examined six turbine based combined cycle Engines. Thermal, chemical, and performance constraints imposed in the propulsion system are also discussed and studied. Genetic optimization technique is employed on each system individually to minimize fuel consumption.
Finally, a comparative study is performed between the six optimized propulsion systems to select the best system to power the indicated aircraft. Performance analysis of the selected propulsion system is presented.

The object of this study is to canvas the literature for the purpose of identifying and compiling a list of Gaps, Obstacles, and Technological Challenges in Hypersonic Applications (GOTCHA). The significance of GOTCHA related deficiencies is discussed along with potential solutions, promising approaches, and feasible remedies that may be considered by engineers in pursuit of next generation hypersonic vehicle designs and optimizations. Based on the synthesis of several modern surveys and public reports, a cohesive list is formed consisting of widely accepted areas needing improvement that fall under several general categories. These include: aerodynamics, propulsion, materials, analytical modeling, CFD modeling, and education in high speed flow physics. New methods and lines of research inquiries are suggested such as the homotopy-based analysis (HAM) for the treatment of strong nonlinearities, the use of improved turbulence models and unstructured grids in numerical simulations, the need for accessible validation data, and the refinement of mission objectives for Hypersonic Air- breathing Propulsion (HAP).

After over 50 years of spaceflight, the calculation of the necessary margin on the thermal protection system of an entering spacecraft remains a largely ad-hoc (in a tactical sense)
process governed by engineering judgment. Over the past several years, NASA has explored a more methodical technique through the application of statistically derived factors to various elements of the design. The current paper first reviews the history of margin application for several prior NASA missions ranging from Apollo to the Mars Exploration Rovers, then discusses a new probabilistic margin process, which was developed for the Mars Science Laboratory and Orion programs. The process is very data-dependent, and
requires validation via a carefully crafted and executed test program including both ground and flight data. Examples of the application of this probabilistic process are given for Mars
Pathfinder, Stardust, and the Mars Science Laboratory. This process is still a work in progress, and the paper discusses some of the key elements of it, along with potential ways
by which the overall methodology could be improved. The eventual goal is the development of a rigorous link between TPS thermal margin process and the corresponding contribution
to the overall heatshield reliability.

The nature of conflict has changed dramatically over the last few decades. Conflicts between States have become the exception rather than the rule. On the other hand, we increasingly witness acute crises and hybrid conflict characterized by internal strife, sometimes in the context of failed or dysfunctional States; or violent separatism, in some cases accompanied by quasi-military operations affecting the civilian population.

The law of war or international (humanitarian) law, i.e. theGeneva and Hague Conventions, is a branch of public international law that sets the acceptable justifications to engage in war (ius ad bellum) and the limits to acceptable wartime conduct (ius in bello). Some principles of customary law became peremptorynorms of jus cogens, a "formal" source of internationallaw that apply erga omnes. Any existing treaty which isin conflict with these norms becomes void and terminates.

A shock tunnel investigation of a scramjet with a rectangular-to-elliptical shape transition (REST) inlet and an elliptical combustor has been conducted at conditions simulating flight at Mach 8.7. The inlet was designed using a quasi-streamline-tracing method to have a design point of Mach 12 and operation down to Mach 6.0. The inlet had a geometric contraction ratio of 6.61, an internal contraction ratio of 2.26, and was preceded by a 150 mm long forebody. The elliptical combustor had both a constant area and diverging section, and was followed by a short thrust nozzle to a total area ratio of 8.0. Gaseous hydrogen fuel was injected via three 4 mm diameter portholes on the intake, and through a series of 1.5 mm diameter portholes on a rearward facing step at the combustor entrance, to promote skin friction reduction as a result of boundary layer combustion. Stable combustion was observed at all fuel equivalence ratios up to 1.23, with positive values of thrust coefficient for the internal flowpath at equivalence ratios above 0.3. These experimental results demonstrate that REST scramjets, designed for access-to-space applications, can operate efficiently at conditions below the design Mach number.

This is a patent application document for my ground to hypersonic Turbine-Ramjet-Scramjet integrated engine (TRS engine) system. The TRS Engine works by applying variable volume combustion chambers and hydraulic based systems to allow for compressor, combustor, and nozzle geometries to be altered during operation. This allows for the integration of a Turbine, Ramjet, and Scramjet all into one system, eliminating the need for a reusable ground to hypersonic vehicle to have any auxiliary propulsion systems onboard.

The supersonic combustion ramjet (scramjet) engine proposed by F.S.Billig in 1964 has been perceived as the most hopeful air breathing propulsion system for the hypersonic flight. The nozzle plays a vital role in accelerating heated air from the combustor to produce thrust in a scramjet engine. In the present study a design methodology has been established using in-house developed python script for the automation process of Computational Fluid Dynamics (CFD), Artificial Neural Network (ANN) and optimization. The nozzle geometry used in this method consist of the inner surface which is represented by a cubic curve. The optimization is carried out for a hypersonic cruise vehicle nozzle designed to fly at Mach number of 6 and an altitude of 30km with single expansion ramp nozzle as the baseline geometry. The results show that the optimized nozzle contour generates 16.7% more thrust when compared to the baseline geometry for similar flow conditions.

It can be argued that at the heart of functional hypersonic vehicle is its engine. Key to a functionally efficient scramjet engine lies in the design of its flow-path. The flow-path is made up of the following sections: (1) the forebody inlet; (2) the isolator, (3) the combustor, and (4) the nozzle. This chapter focuses on the design of the forebody inlet and the isolator sections of a scramjet engine. In this framework, key to a functionally efficient scramjet engine lies in the design of its flow-path. This flow-path design must consider a complex flow-field physics and the interaction of physical surfaces with this complex flow-field. Many attempts to design efficient scramjet flow-paths have met with some measured degree of success. This research uses a 'inverse design' approach, which is similar to Darwin's theory of evolution, where an organism adopts to survive in its environment; the scramjet flow-path will be carved/extracted from the operational environment. The objective is to naturally and organically capture, process and direct the flow from the environment; thus preparing it for the combustion process. This approach uses the ideal 2-D oblique shock relations, coupled with Nonweiler's caret waverider theory and streamline marching techniques.

A = decomposing rate coefficient, 1∕s B 0 = normalized ablation rate, _ m c ∕ρ e U e C H C p = specific heat, J∕kg ⋅ K E = activation energy, J∕mol h = enthalpy, J∕kg h = mass weighted averages of the material enthalpy, J∕kg _ m g = pyrolysis gas mass flux, kg∕m 2 ⋅ s k = thermal conductivity, W∕m ⋅ K R = universal gas constant, 8.314 J∕mol ⋅ K _ s = surface recession rate, m∕s T = temperature, K t = time, s y = in-depth direction coordinate, m ε = emissivity coefficient ρ = density, kg∕m 3 ρ e U e C H = heat transfer coefficient, kg∕m 2 ⋅ s σ = Stefan-Boltzmann constant, W∕m 2 ⋅ s ⋅ K 4 ψ = reaction order Subscripts c = char e = freestream g = pyrolysis gas r = recovery s = surface v = virgin w = wall ∞ = background condition

Sparta is a three-dimensional finite volume compressible Java based flow solver on multiple block structured grids. The flow solver is graphical user interface driven and platform independent as well. Finite-rate reaction kinetics, thermal and chemical non-equilibrium and high temperature transport coefficients are all developed and incorporated into the solver. Zero, one and two equation turbulence models have been implemented in the solver for turbulence modeling. The flow domain is discretized by a structured grid and a finite-volume approach is used to discretize the conservation equations. One of several schemes can be chosen from the user interface to calculate inviscid fluxes across cell faces while central differences schemes are used to calculate the viscous fluxes. The CFD solver includes a geometry engine to generate surface and volume grids, a trajectory engine to generate trajectory data given the appropriate atmospheric models, an aerodynamic flow solver to calculate the aerodynamic forces and moments and finally hooked to the Navier-Stokes engine. Probe models can also be selected automatically from the graphical user interface and the database. The database comprises vehicle dimensions, trajectory data, and aero-thermal, Thermal Protection Systems data for many different ballistic entry vehicles. Material properties for carbon and silicon based ablators have been obtained and can be accessed from the database. Comparative data analysis capability is achieved through a Relational Database Management System. Atmospheric models can be selected from the graphical user interface so that the multi-species to be solved are automatically populated before the solver is invoked. Sparta has been integrated to a planetary probe database and a trajectory code to study trajectory analysis, aerodynamic heating and flow-field analysis. The purpose of this research is twofold: to investigate a pure Java based multi-threaded CFD solver and to implement a semi-automatic and interactive solver for aerothermodynamics such that grids can be generated automatically given the geometry of the vehicles, trajectory data can be generated given the atmospheric model and TPS material from the graphical interface for the generation of flow field information for atmospheric entry vehicles. Sparta accepts hooks from the trajectory so that inputs for the flow solver can come from the trajectory output making the solver trajectory driven. This paper addresses the CFD development, implementation and capabilities of a hypersonic flow solver Sparta.

Keywords: aerothermodynamics; Navier-Stokes Solutions; TPS; hypersonics; automatic grid generation; reentry; databases; framework

Hypersonic weapons are becoming a reality with the deployment of Avangard glide vehicle by Russia and the successful testing of such weapons by China and the US. Built to overcome ballistic missile defenses, these weapons are often touted as unstoppable and a possible game-changer. Unlike traditional missiles, hypersonic glide vehicles glide towards their targets giving the defender no time to respond and intercept them. These weapons impact strategic stability by compressing the decision-making time, increasing counterforce temptations due to their precision targeting, and by causing a new arms race. The ability to carry both conventional and nuclear warheads off these weapon can also lead to miscalculation and inadvertent escalation. With weakening arms control and trust deficit at the global level, the global strategic stability would be under stress with the induction of these weapons in nuclear-weapon states' inventories until some confidence-building measures are adopted. They will continue to challenge the strategic stability even after new systems are developed to counter them.

Radio blackout is commonly associated with interference from plasma created around high-speed aerospace and re-entry vehicles. Alternative communication means, including lasers at optical frequencies, has emerged as promising counter to the problem. The refractive index of a medium, such as air, governs the angular shift in the path of an optical signal. For a fluid, the refractive index is a function of the thermodynamic state. Drastic changes in the thermodynamic variables across the shock wave changes the downstream refractive index. This paper presents an investigation of how changes to the thermodynamic state of air for high-speed flow (not including chemical reactions) affects optical propagation. Analytical expressions for horizontal deviation and angular shifts induced by-shock layer, shock wave and atmosphere, are derived for an optical signal that travels from a high-speed vehicle to the ground. The formulation has been verified computationally for the supersonic flow about a wedge and cone that capture effects from flow gradients.

UNSW Canberra has been running experiments to measure the aerodynamics coefficients and the external flow fields on free-flight models of high speed aerial vehicles in ground-based test facilities. It is found that the dynamic behaviour of the free flying models is a function of both the aerodynamics of the flow over the model and also the mass distribution of the model. A clear analytical model of the scaling of the dynamic behaviour is vital to both make sense of the measured dynamic response of the model and to select an optimum mass distribution of the model for an experiment. The shuttle orbiter was used as a candidate geometry for the scaling laws as it travels at hypersonic speeds. It was found that the scaling laws are in good agreement with real flight data. The ratio of air density between shuttle altitude to that at model altitude was an important factor when scaling the mass of the model. This affected the mass of the model and how the wind tunnel conditions are set. The altitude at which the shuttle orbiter experiments are carried out affected the sensitivity of the air density ratio.

A report of the M12REST scramjet ground test program
at ‘on
-
design’ test
conditions, conducted from January 1
st
to June 24
th
, 2009 in the T4 Shock Tunnel Facility at the Centre for Hypersonics, The University of Queensland, is presented. The study was performed to investigate discrepancies between numerical and experimental results of a previous 2007 test program involving the M12REST scramjet engine. Off-design results of the engine in the 2007 study demonstrated good agreement between numerical and experimental results (Suraweera and Smart, 2009). However, on-design experimental results showed a large pressure region on the forward section of the inlet that could not be replicated using a computational fluid dynamics (CFD) code (White and Morrison, 1999). The present study trialled combinations of ten distinct boundary layer trip configurations, in order to investigate whether this large pressure region was the result of local flow separation. A blunt 3 mm radius leading edge and a longer 500 mm forebody were also separately tested.
Three new ‘on
-
design’ flow conditions
(four in total) were also tested. Pressure and heat transfer measurements were taken along the engine flowpath.  A description of the T4 Shock Tunnel and its operating characteristics has been given. Drawings of the proposed test model and extraneous test articles have also been provided. All 45 test runs executed during the experimental campaign have been listed, along with the corresponding flow properties for four test conditions. Mean pressure and Stanton number distributions for significant tunnel runs, illustrating the effects of various boundary layer trip and engine fuelling configurations, have been presented. The fuel used was gaseous hydrogen. Supersonic and subsonic combustion was measured at a range of fuel equivalence ratios for two of the test conditions (3 and 4). Inlet injection was found to produce separated flow regions within the inlet, and hence the fuelling scheme was discontinued. Step injection was tested successfully at a range of equivalence


3
ratios. In terms of combustion induced increases in pressure, higher levels were seen when engine was run with the M11 enthalpy test condition 4. However, the engine was able to operate in true scramjet mode with the inflow of the M12 enthalpy test condition 3. Furthermore, the engine was found to be more stable at this condition as the combustion induced pressure rise was contained by the isolator.

A shock tunnel investigation of a scramjet with a rectangular-to-elliptical shape transition (REST) inlet and an elliptical combustor has been conducted at conditions simulating flight at Mach 8.7. The inlet was designed using a quasi-streamline-tracing method to have a design point of Mach 12.0 and operation down to Mach 6.0. The elliptical combustor began with a rearward facing step around its perimeter, and was followed by a constant area and diverging section. The flowpath was completed by a short thrust nozzle. Gaseous hydrogen fuel was injected either through multiple portholes on the intake, or a series of 48 portholes on the rearward facing step at the combustor entrance, or a combination of the two. All fuelling configurations resulted in a positive thrust coefficient at equivalence ratios above 0.3 without the use of ignition aids. Fuel injection in the intake produced robust combustion and good internal thrust levels, but led to inlet unstart at fuel equivalence ratios above 0.61. Stable, mixing limited combustion was observed for fuel injection at the step at all fuel equivalence ratios up to 1.23. Combined intake and step injection was observed to have the best performance. These experimental results demonstrate that REST scramjets, designed for access-to-space applications, can operate efficiently at conditions below the design Mach number.

his dissertation examines the e↵ects of density gradients on turbulent boundary layer statistics and structure using Particle Image Velocimetry (PIV). Two distinct cases were examined: the thermally stable atmospheric surface layer characteristic of nocturnal or po- lar conditions, and the hypersonic bounder layer characteristic of high speed aircraft and reentering spacecraft.
Previous experimental studies examining the e↵ects of stability on turbulent boundary layers identified two regimes, weak and strong stability, separated by a critical bulk stratifi- cation with a collapse of near-wall turbulence thought to be intrinsic to the strongly stable regime. To examine the characteristics of these two regimes, PIV measurements were ob- tained in conjunction with the mean temperature profile in a low Reynolds number facility over smooth and rough surfaces. The turbulent stresses were found to scale with the wall shear stress in the weakly stable regime prior relaminarization at a critical stratification. Changes in profile shape were shown to correlate with the local stratification profile, and as a result, the collapse of near-wall turbulence is not intrinsic to the strongly stable regime. The critical bulk stratification was found to be sensitive to surface roughness and potentially Reynolds number, and not constant as previously thought. Further investigations examined turbulent boundary layer structure and changes to the motions that contribute to turbulent production.
To study the characteristics of a hypersonic turbulent boundary layer at Mach 8, sig- nificant improvements were required to the implementation and error characterization of PIV. Limited resolution or dynamic range e↵ects were minimized and the e↵ects of high shear on cross-correlation routines were examined. Significantly, an examination of parti- cle dynamics, subject to fluid inertia,compressibility and non-continuum e↵ects, revealed that particle frequency responses to turbulence can be up to an order of magnitude smaller than estimates made using a standard shock response test. The e↵ect of over-large tripping
devices was also found to increase the wake strength of the mean velocity profile as well as freestream turbulence. A final assessment of the data reveals that Morkovin scaling col- lapses the streamwise turbulence profiles with DNS at the same Mach number. Wall-normal turbulence measurements remain compromised by limited particle frequency response.

Military emerging disruptive technologies (EDTs) have a rapid and major effect on technologies that already exist and disrupt or overturn traditional practices and may revolutionize governmental structures, economies and international security. Margaret Kosal (2019) finds that military applications of EDT have even greater potential than nuclear weapons to radically change the balance of power. The debate, stimulated by the Group of Governmental Experts on emerging technologies in the area of lethal autonomous weapons systems (GGE on LAWS) established by the Convention on Prohibitions or Restrictions on the Use of Certain Conventional Weapons (CCW), has focused on artificial intelligence, cyber weapons and drones. A broader discussion should include all military EDTs, inter alia: space and hypersonic weapons; directed-energy weapons/laser and photonic weapons, just to name a few. Though leaders have begun to become aware of legal and ethical implications of the military use of EDTs, these issues remain in the background: security concerns are of pivotal importance and a most of the information and documents are kept confidential and their circulation is restricted. This paper aims to investigate the compliance of military EDTs with international law, international humanitarian law, international human rights law and with ethical principles.

As weapons are getting more sophisticated, researchers are trying to build defense weapon systems that can keep pace with the continuously evolving technology. Just as stealth is changing the nature of warfare today, speed is being seen as the next big advancement in the field of aviation. After attaining stability in supersonic systems scientists are aiming towards the next segment of speed i.e. hypersonic region where speed ranges above Mach 5. The United States Air Force Chief Scientist Mica Endsley announced the target date for the next generation Hypersonic aircraft as 2023 . This time it will take the form of a 6500 kilometers per hour unmanned strike aircraft known as SR-72. It is by far the most advanced and funded hypersonic aircraft project the country has till date, hence the most relevant candidate for the claims by the US Air Force. Countries such as China and Russian are also developing in hypersonic technology because of non-availability of weapons to counter act these systems. The need for hypersonic weapons has emerged because with the evolving weapon technology which are intended to strike aircrafts, speed is the only alternative to escape the influence region of such weapons. Hypersonic aviation does not only provide high speed but along with it the capability to cover great distances at that speed .

Nomenclature m non-dimensional mass injection rate qT, cold total cold-wall heating rate, including both convection and radiation, W/m 2 U∞ free stream velocity, m/s ρ∞ free stream density, kg/m 3 R ratio of radiative heating rate corrected for blowing effect to "cold wall," radiative heating rate

Results from an experimental and numerical study of skin friction levels obtained when hydrogen is injected into turbulent boundary layers are presented. Measurements are reported from experiments in the T4 free-piston reflected shock tunnel. Hydrogen was injected from a 3 mm high slot into the boundary layer on the flat surface of one of the walls of a duct 100 mm wide, 60 mm high, and 1745 mm long. The experiments were conducted at Mach numbers ranging from 4.2 to 4.7, flow stagnation enthalpies of 4.8 MJ/kg to 9.5 MJ/kg, static pressures of 59 kPa to 86 kPa, and Reynolds numbers of 8.9  106 m-1 to 17.2  106 m-1. Hydrogen fuel was injected 245 mm downstream of the inlet through a 3 mm slot. Mass flow rates of 0 kg/s/m and 0.36 kg/s/m, at a nozzle area ratio of 1.7, were used for test flows of air. Combustion occurred at all flow conditions with results indicating a maximum reduction in skin friction coefficient, of approximately 80% of the level measured with no injection. Skin friction reductions of approximately 60% were obtained at two other test flows. Measured heat transfer levels were found to be comparable with levels obtained without injection, for most of the experimental conditions. Hydrogen injection into a test flow of nitrogen was also trialed at all flow conditions to compare with the results obtained when fuel was injected into an air flow, in order to identify the effects of combustion. In general, the results showed that reductions in local skin friction coefficient were greater when combustion occurred than when fuel was injected and did not burn.  This paper is unpublished.

Results from an experimental and numerical study of skin friction levels obtained when hydrogen is injected into turbulent boundary layers are presented. Measurements are reported from experiments in the T4 free-piston reflected shock tunnel. Hydrogen was injected from a 3 mm high slot into the boundary layer on the flat surface of one of the walls of a duct 100 mm wide, 60 mm high, and 1745 mm long. The experiments were conducted at Mach numbers ranging from 4.0 to 4.5, flow stagnation enthalpies of 4.8 MJ/kg to 9.5 MJ/kg, static pressures of 75 kPa to 110 kPa, and fuel equivalence ratios of 0.3 and 0 for test flows of air. Combustion occurred at all flow conditions with results indicating a maximum reduction in skin friction coefficient, of approximately 77% of the level measured with no injection. Skin friction reductions of approximately 60% were obtained at two other test flows. Measured heat transfer levels were found to be comparable with levels obtained without injection, for most of the experimental conditions. Hydrogen injection into a test flow of nitrogen was also trialed at all flow conditions to compare with the results obtained when fuel was injected into an air flow in order to identify the effects of combustion.

Die Corona Chronik arbeitet das Verhältnis von Corona Lockdowns, Social Media, Klimawandel, Internetsucht, Künstliche Intelligenz, Digitalisierung und Bewegungsmangel auf.

Dr. Pawlik folgt in seiner Analyse McLuhan, Heinz von Foerster und Chuangtse.

The effect of varying three-dimensional, cylindrical post-type trip size on the mean and turbulent velocity profiles of a Mach 7.6 turbulent boundary layer is examined using particle image velocimetry. It is shown that the effect of under-and overtripping is to amplify the wake component of the mean velocity profile and outer-layer turbulence intensity, confirming trends from incompressible flow. Such results indicate that overly aggressive tripping introduces artificial large-scale turbulence that requires longer downstream distances to decay. For the current experiment, adequate tripping was obtained for trip sizes between 1.7 and 2.3 times the laminar boundary-layer displacement thickness at the trip, δ tr , with the optimum size approximately 2.3 δ tr. The wake strength for the optimally tripped cases followed the correlation of Fernholz and Finley (AGARDograph 253, Neuilly sur Seine, France, 1980) at the same Reynolds number, providing a good indicator for under- or overtripping. These results confirm that compressible boundary layers mimic incompressible trends but require larger trip sizes and have increased sensitivity, making a boundary layer free from initial conditions harder to achieve.

The mechanism of Tunguska blast is determined as quick destruction of four main fragments of the comet in dense atmosphere. The structure of comet is determined, its substance and final size of each fragment. Detailed explanation of thermal damages due to explosion is described. More accurate coordinates of regions are obtained on the area of Tunguska site, where the most heavy sediments or fragments are possible to discover. Quasi three-dimensional modeling of the impact was produced with the help of PIC-method. It was shown that Tunguska-size comets are able to penetrate considerable deep into dense atmosphere due to decrease of drag effect. This decrease is explained by forward-directed jet from cavern of ice body, which located at the region of stagnation point. There is additional information for emergency organizations for training of people against comets and asteroids danger, its mechanical, thermal and electromagnetic influence.

HIFiRE 7 Flight Test Update 29 April, 2015.

Recently, the HIFiRE 7 flight was launched from the western coast of Norway.  The HIFiRE 7 payload consisted of two back-to-back three-dimensional scramjet engines set within a trailing flare for aerodynamic stability during deployment as a free-flyer.  The scientific goals of the flight were to perform a direct measurement of the thrust generated by the scramjet flowpaths, and to compare the measured thrust with estimates based on ground testing.

To date, it was one of the most advanced scramjet flowpaths ever flight tested.  The launch and most of the return leg of the flight performed as planned.  However, no telemetry and hence no engine data was returned to ground stations during the experimental window of the flight envelope.  Presently, the full characterisation of the failure is undetermined.  The payload and the remaining stack impacted the surface within operational guidelines.

The flight represented years of design and testing by a relatively small group of extremely, hardworking people.  I thank you all for your efforts.  Shit happens…

M.Suraweera

Russia's testing of an SSC-X-9 Skyfall hypersonic weapon has rekindled interest in a largely forgotten Cold War technology: nuclear-powered air power. In this article, Peter Layton looks at the history of US development of nuclear-powered weapons in the 1950s and 1960s. In doing so, he highlights that successful innovation does not necessarily require a system to be operationally fielded.

High-fidelity hypersonic computations have become practical in recent years. Large Eddy Simulations (LES) and high order shock-capturing schemes, such as ENO and WENO, are able to compute high-speed, three-dimensional flow. However, common application of such schemes is limited by high computational costs and uncertainty errors that are difficult to quantify. A novel technique called Integro-Differential Scheme (IDS) may have the capability to alleviate these current limitations. Unlike common numerical techniques where the solution is integrated in time using local information, this method reconstructs the temporal and spatial fluxes using a consistent averaging procedure. This paper focuses on the accuracy with which the IDS predicts the complex physics involved in hypersonic flow devices such as isolators and flame holders. Accurate numerical prediction of such devices is critical in the design and development of supersonic combustion air-breathing engines, also known as Scramjets. The goal of the study is to demonstrate the validity of the IDS by computing supersonic flow in an isolator and flame holding cavity for comparison to experimental data.

Flight pressure and heat flux data have been compared to angle-of-attack-and yaw-dependent computational-fluid-dynamics results for pressure distribution as well as laminar and turbulent heat-transfer results at three time points in the ascent trajectory. Computed pressures, normalized by freestream pressure, were interpolated to the flight Mach numbers at each time point throughout the ascent and descent trajectories, and angle of attack and yaw were estimated from measured pressure by determining the combination minimizing the difference between the measured and computed pressures. The resulting vehicle attitude was compared to the vehicle attitude derived from inertial measurement unit results from the flight. The two methods showed excellent agreement for the entirety of the ascent and reentry portions of the trajectory. A similar normalization of the laminar and turbulent computational-fluid-dynamics heat transfer results into Stanton number distributions was compared to flight heat transfer measurements, and transition times at a given location were inferred. Computational heat conduction analysis verified assumptions in the calculation of heat flux from temperature. Nomenclature c p = specific heat capacity, constant pressure, J∕…kg ⋅ k† M = Mach number p = pressure, Pa _ q = heat flux, W∕m 2 Re = Reynolds number St = Stanton number T = temperature, K t = time, s u = velocity, m∕s x = streamwise distance, m α = angle of attack, deg β = yaw, deg ρ = density, kg∕m 3 ϕ = angular location on vehicle surface, deg Subscripts CFD = derived from computational solution CL = centerline e = boundary-layer edge conditions F = derived from flight measurement FFT = fast Fourier transform IMU = recorded by the inertial measurement unit LE = leading edge RMS = derived from root-mean-squared error minimization between computational fluid dynamics and flight tr = at transition onset w = wall conditions 0 = stagnation conditions ∞ = freestream conditions

A novel method to delay transition in hypervelocity flows in air over slender bodies by injecting CO 2 into the boundary layer is presented. The dominant transition mechanism in hypersonic flow is the inviscid second (Mack) mode, which is associated with acoustic disturbances which are trapped and amplified inside the boundary layer . In dissociated CO 2 -rich flows, nonequilibrium molecular vibration damps the acoustic instability, and for the high-temperature, high-pressure conditions associated with hypervelocity flows, the effect is most pronounced in the frequency bands amplified by the second mode . Experimental data were obtained in Caltech's T5 reflected shock tunnel. The experimental model was a 5 degree half-angle sharp cone instrumented with 80 thermocouples, providing heat transfer measurements from which transition locations were from turbulent intermittency based upon laminar and turbulent heat flux correlations. An appropriate injector was designed and fabricated, and the efficacy of injecting CO 2 in delaying transition was gauged at various mass flow rates, and compared with both no injection and chemically inert Argon injection cases. Argon was chosen for its similar density to CO 2 . At an enthalpy of approximately 10 MJ/kg (Eckert's reference temperature T * = 2550 K), transition delays in terms of Reynolds number were documented. For Argon injection cases at similar mass flow rates, transition is promoted.

The results of a seminal experimental study of the effects of bluntness and swallowing length on transition on an 8-degree cone at zero angle of attack in Mach 6 high Reynolds number flow are analyzed with the STABL computational fluid dynamics code package. Mean flow solutions and PSE-Chem stability analyses for a total of 11 different nose tip bluntnesses, ranging from sharp to 15.24 mm radius, are obtained. For the sharpest cases, NTr ≈ 7, but as bluntness increases and the calculated swallowing distance lengthens, the computed N-factor at the experimentally-observed transition location drops below the level at which Mack’s second mode would be expected to lead to transition. These results indicate that the dominant instability mechanism for the bluntest cases is not the second mode.

A high-fidelity simulation of the massively separated shock/transitional boundary layer interaction caused by a 15-degrees axisymmetrical compression ramp is performed at a free stream Mach number of 6 and a transitional Reynolds number. The chosen configuration yields a strongly multiscale dynamics of the flow as the separated region oscillates at low-frequency, and high-frequency transitional instabilities are triggered by the injection of a generic noise at the inlet of the simulation. The simulation is post-processed using Proper Orthogonal Decomposition to extract the large scale low-frequency dynamics of the recirculation region. The bubble dynamics from the simulation is then compared to the results of a global linear stability analysis about the mean flow. A critical interpretation of the eigenspectrum of the linearized Navier-Stokes operator is presented. The recirculation region dynamics is found to be dominated by two coexisting modes, a quasi-steady one that expresses itself mainly in the reattachment region and that is caused by the interaction of two self-sustained instabilities, and an unsteady one linked with the separation shock-wave and the mixing layer. The unsteady mode is driven by a feedback loop in the recirculation region, which may also be relevant for other unsteady shock-motion already documented for shock-wave/turbulent boundary layer interaction. The impact of the large-scale dynamics on the transitional one is then assessed through the numerical filtering of those low wavenumber modes; they are found to have no impact on the transitional dynamics.

A series of slender-body hypervelocity boundary-layer instability and transition experiments were performed in the Caltech T5 Reflected Shock Tunnel. During this campaign, it became clear that the condition of the T5 shock tube would significantly affect the consistency of the instability and transition measurements; a regimen of cleaning was iterated on until satisfactory repeatability was achieved. In this work, a description of the cleaning regimen is given. Additionally, boundary-layer instability measurements and a statistical analysis of the boundary-layer transition scatter are presented for experiments before and after cleaning regimen implementation.

SUMMARY Turbulence modeling for high-speed compressible flows is described and discussed. Starting with the compressible Navier-Stokes equations, methods of statistical averaging are described by means of which the Reynolds-averaged Navier-Stokes equations are ...