Constant Dynamic Pressure Research Articles

Multipoint Design Optimization of Busemann-Based Intakes for Scramjet-Powered Ascent Flight

Scramjet engines are one of the most promising hypersonic airbreathing propulsion technologies for robust, efficient, and economical access to space. Multi-objective design optimization has been conducted for Busemann-based intakes in inviscid and viscous regimes at single and multiple design points, respectively, by means of surrogate-assisted evolutionary algorithms coupled with computational fluid dynamics in the present research. Intake geometries are generated by applying geometric alterations to the full Busemann intake via leading-edge truncation, axial stunting, and radial contraction, aiming to simultaneously minimize intake drag and maximize the compression efficiency at two different design conditions (i.e., Mach 7.7 at an altitude of 30 km and Mach 10 at 33.5 km on a constant dynamic pressure ascent trajectory). The single-point inviscid optimization study has found that the nondominated solutions obtained from minimizing drag and maximizing compression efficiency are approximately the same as those obtained from maximizing total pressure recovery and minimizing static pressure ratio. From the multipoint viscous optimization study, the optimal solutions have been found to retain the original advantages of the inviscid full Busemann intakes in terms of high compression efficiency with shorter intakes and higher static pressure as well as adequately high mean exit temperature for both design conditions.

Numerical investigation of axisymmetric intake flowfield and performance for scramjet-powered ascent flight

Air intakes play a crucial role in the operation of scramjet engines, one of the most promising hypersonic airbreathing propulsion technologies for efficient and flexible space transportation, it is thus essential to accurately understand and predict the flowfields for high-performance intake design, especially in varying operating conditions. Axisymmetric intake flowfields are investigated numerically assuming a typical scramjet operation on a constant dynamic pressure ascent trajectory, aiming to achieve scramjet-powered access-to-space. Physical insight into scramjet flowfields at various flight altitudes and Mach numbers have been gained by means of high-resolution simulation to resolve the viscous flowfields. High levels of adaptive mesh refinement have identified detailed flow phenomena including shock reflection, interactions, and boundary layer as well as their aerodynamic and aerothermal effects on pressure and temperature fields. In particular, Mach reflection characterized by a Mach disk has been found to occur at the centerline, even in the operating conditions where regular reflection would result from computational simulation without mesh adaption. The intake performance has been evaluated with respect to the compression efficiency, drag, and exit temperature. The intake drag has been found to decrease at higher Mach numbers at higher altitudes on a constant dynamic pressure trajectory during acceleration due to reduced pressure drag while viscous drag varies little. Rather little influence of varying wall conditions has been observed on the intake drag between the isothermal and adiabatic wall conditions due to counteracting behavior between viscous and pressure drag components.

Development and Application of Computational Tool Using Local Surface Inclination Methods for Preliminary Analysis of Hypersonic Vehicles

This work presents a computational tool for preliminary analysis of hypersonic vehicles, based on local surface inclination methods: the HipeX. This program was developed for reading standard triangulation language (STL) geometry files and calculating pressure coefficient and temperature distributions over vehicle’s surface using the Newtonian, modified Newtonian or tangent-wedge methods. Validations were made with a cylinder and a sphere, where only the Newtonian method was applied, and with experimental data from Apollo capsule at Mach 10, where the Newtonian and the modified Newtonian methods were applied. These validations presented the code capability to read geometries as well as to estimate aerodynamic force coefficients. A preliminary application was to predict the aerodynamic force coefficients of a generic hypersonic vehicle over constant dynamic pressure trajectories of 23,940, 60,000 and 95,760 N/m2 with zero angle of attack. With a fixed dynamic pressure of 60,000 N/m2, this vehicle was tested over several Mach numbers and with angle of attack variation from -10 to 10 deg. Zero angle of attack investigation showed minor changes on the force coefficients with altitude, while the variation of angle of attack produced more pronounced variations on these parameters. Maximum flow temperatures over vehicle’s surface were estimated ranging from 850 to 5,315 K.

Pulse wave response characteristics for thickness and hardness of the cover layer in pulse sensors to measure radial artery pulse

BackgroundPiezo-resistive pressure sensors are widely used for measuring pulse waves of the radial artery. Pulse sensors are generally fabricated with a cover layer because pressure sensors without a cover layer are fragile when they come into direct contact with the skin near the radial artery. However, no study has evaluated the dynamic pulse wave response of pulse sensors depending on the thickness and hardness of the cover layer. This study analyzed the dynamic pulse wave response according to the thickness and hardness of the cover layer and suggests an appropriate thickness and hardness for the design of pulse sensors with semiconductor device-based pressure sensors.MethodsPulse sensors with 6 different cover layers with various thicknesses (0.8 mm, 1 mm, 2 mm) and hardnesses (Shore type A; 30, 43, 49, 71) were fabricated. Experiments for evaluating the dynamic pulse responses of the fabricated sensors were performed using a pulse simulator to transmit the same pulse wave to each of the sensors. To evaluate the dynamic responses of the fabricated pulse sensors, experiments with the pulse sensors were conducted using a simulator that artificially generated a constant pulse wave. The pulse wave simulator consisted of a motorized cam device that generated the artificial radial pulse waveform by adjusting the stroke of the cylindrical air pump and an air tube that conveyed the pulse to the artificial wrist.ResultsThe amplitude of the measured pulse pressure decreased with increasing thickness and hardness of the cover layer. Normalized waveform analysis showed that the thickness rather than the hardness of the cover layer contributed more to waveform distortion. Analysis of the channel distribution of the pulse sensor with respect to the applied constant dynamic pressure showed that the material of the cover layer had a large effect.ConclusionsIn this study, in-line array pulse sensors with various cover layers were fabricated, the dynamic pulse wave responses according to the thickness and the hardness of the cover layer were analyzed, and an appropriate thickness and hardness for the cover layer were suggested. The dynamic pulse wave responses of pulse sensors revealed in this study will contribute to the fabrication of improved pulse sensors and pulse wave analyses.

Scramjet Performance with Nonuniform Flow and Swept Nozzles

Numerical simulations are presented that compare the specific impulse of a generic inlet-fueled uniform-compression and nonuniform-compression scramjet over the flight Mach number range of 7–12 with a constant freestream dynamic pressure trajectory of 100 kPa. Both engine configurations have identical inlet contraction ratio, inlet to nozzle area ratio, and combustor and nozzle geometries. The results show that the specific impulse in the nonuniform-compression scramjet is higher than the uniform-compression scramjet at each flight condition. The increase in specific impulse for Mach 8, 9, 10, 11, and 12 are 165, 178, 96, 74, and 15 s, respectively. Both the uniform- and nonuniform-compression engines unstarted at the Mach 7 condition. In addition, two three-dimensional swept-nozzle configurations (45 and 60 deg) are integrated into the nonuniform-compression engine to investigate an idea of coupling the nonuniform flow with a nonuniform expansion nozzle to improve performance. The 45 deg results show no change in specific impulse at Mach 8–11 and a 50 s improvement at Mach 12. The 60 deg configuration has less specific impulse at each flight condition, except Mach 12. The thermal loads are reduced by 15 and 30% for the 45 and 60 deg configurations at each flight condition. The findings provide new insight into how nonuniform compression with swept nozzles can provide an additional degree of freedom in the design of fixed-geometry scramjets that have robust performance over a range of flight conditions.

Effect of Dynamic Pressure on the Performance of Thermoacoustic Refrigerator with Aluminium (Al) Resonator

In practice the refrigerants are being used in the conventional refrigeration system to get the required cooling effect. These refrigerants produce Chlorofluorocarbons (CFCs) and Hydro chlorofluorocarbons (HCFCs) which are highly harmful to the environment, particularly depleting of ozone layers resulting in green house emissions. In order to overcome these effects, the research needs to be focused on the development of an ecofriendly refrigeration system. The thermoacoustic refrigeration system is one among such system where the sound waves are used to compress and expand the gas particles. This study focuses on the effect of dynamic pressure on the thermoacoustic refrigerator made of aluminium with overall length of 748.82 mm, and the entire inner surface of the resonator tube was coated with 2mm thickness of polyurethane to minimize the heat losses to the atmosphere. Experiments were conducted with different stack geometries i.e. parallel plates having 0.119 mm thick with spacing between the plates maintained at 0.358 mm, 1mm diameter pipes, 2mm diameter pipes and 4 mm diameter pipes. Experiments were also conducted with different drive ratios of 0.6%, 1% and 1.6% for a constant dynamic pressure of 2 bar and 10 bar for helium and air as working medium. The results were plotted with the help of graphs, the variation of coefficient of performance (COP) and the relative coefficient of performance (COPR) for the above said conditions were calculated.

Ascent trajectory tracking for an air-breathing hypersonic vehicle with guardian maps

The emphasis of this article lies in the development of a gain scheduling algorithm for the constant dynamic pressure ascent trajectory. Due to the broad flight envelope and nonlinear characteristics of hypersonic vehicles, a design method of adaptive control based on guardian maps is proposed. Two types of ascent trajectory are considered: acceleration within Scram mode and acceleration across the Ram–Scram mode transition. A velocity-based linear parameter varying model of an air-breathing hypersonic vehicle is established. The gain vector of controller is determined via optimal control theory based on linear quadratic regulator technology with the specified flying qualities. With the iteration of proposed guardian maps approach, the set of controller gain vector within the flight envelope is obtained. The results of nonlinear simulation show that the proposed approach of control system design has the ability to ensure the global stability through the flight envelope and to maintain the flying qualities.

Automatic Approach and Landing Trajectory Planner for Unpowered Reusable Launch Vehicle

A new guidance scheme for the approach and landing (A & L) phase of an unpowered reusable launch vehicle (RLV) has been developed. The main advantage of the new guidance is the use of glide-efficiency factor as the independent variable to compute the geometrical flare parameters by a set of analytical functions. The trajectory-planning algorithm generates its reference geometry based on the steep and shallow subphases, respectively. During the steep segment, the quasi-equilibrium glide (QEG) solution, which assumes a constant dynamic pressure and flight-path angle during the flight, is used to create the flight reference while the shallow segment is defined by polynomial functions for altitude and dynamic pressure profiles. Standard linearization methods are used to design a closed-loop command in order to track the QEG profile. Furthermore, proportion-derivative (PD) control is used to modulate the lift coefficient during the flare flight. Once the reference trajectory is created, a closed-loop simulation is obtained to track the reference. Off-nominal conditions, in terms of change in initial glide-efficiency factor, dynamic pressure, flight-path angle, and altitude are tested using a Monte-Carlo simulation. The simulated results demonstrate the effectiveness of the proposed algorithm to land the vehicle successfully under large dispersions of glide-efficiency factor.

Aerodynamic Shape Design of Hypersonic Booster with RBCC Engine for TSTO Vehicle

An appropriate configuration of a hypersonic booster plane with the rocket-based-combined-cycle (RBCC) engine was studied for the two-stage-to-orbit (TSTO) reusable space transportation system. Assuming the flight path at a constant dynamic pressure followed by a constant acceleration G, the ascent trajectory and the vehicle sizing were numerically analyzed. The feasibility of the design was evaluated by the ratio of the vehicle mass after the orbiter separation to the takeoff mass. A rocket-propelled orbiter of 75 ton with the separation speed at 3000 m/s was considered as the upper stage. The results show that a vehicle with the takeoff mass smaller than 600 ton including the orbiter can be realized under the constraint of the maximum acceleration at 1 G. The reduction in the drag force to the level of CD=0.01 is strictly required for successful TSTO design. A vehicle with the RBCC engine can continue to accelerate by using the rocket mode in low atmospheric density environment, and has an advantage to conduct the orbiter separation at a high altitude, avoiding hazardous aerodynamic interference. From a hypersonic aerodynamic viewpoint, a waverider-derived vehicle with twin engines on the upper surface and the internal bay, in which the orbiter is placed and released backward, is recommended.

Internally driven large-scale changes in the size of Saturn's magnetosphere.

Saturn's magnetic field acts as an obstacle to solar wind flow, deflecting plasma around the planet and forming a cavity known as the magnetosphere. The magnetopause defines the boundary between the planetary and solar dominated regimes, and so is strongly influenced by the variable nature of pressure sources both outside and within. Following from Pilkington et al. (2014), crossings of the magnetopause are identified using 7 years of magnetic field and particle data from the Cassini spacecraft and providing unprecedented spatial coverage of the magnetopause boundary. These observations reveal a dynamical interaction where, in addition to the external influence of the solar wind dynamic pressure, internal drivers, and hot plasma dynamics in particular can take almost complete control of the system's dayside shape and size, essentially defying the solar wind conditions. The magnetopause can move by up to 10–15 planetary radii at constant solar wind dynamic pressure, corresponding to relatively “plasma‐loaded” or “plasma‐depleted” states, defined in terms of the internal suprathermal plasma pressure.

An Approximate Optimal Maximum Range Guidance Scheme for Subsonic Unpowered Gliding Vehicles

This study investigates the maximum gliding range problems of subsonic unpowered gliding vehicles and proposes an approximate optimal maximum range guidance scheme. First, the gliding flight path angle corresponding to constant dynamic pressure is derived. A lift-to-drag ratio(L/D)inversely proportional to the dynamic pressure is then proven. On this basis, the calculation method of an optimal dynamic pressure (ODP) profile with a maximumL/Dthroughout the flight is presented. A guidance scheme for tracking the ODP profile, which uses the flight path angle as control variable, is then designed. The maximum ranges of the unpowered gliding vehicle obtained by the proposed guidance scheme and pseudospectral method are compared. Results show that the guidance scheme provides an accurate approximation of the optimal results, and the errors are less than 2%. The proposed guidance scheme is easy to implement and is not influenced by wind compared with numerical schemes.

Mach Hydrocarbon-Fueled Scramjet Flight Experiment: The HIFiRE Flight 2 Project

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a collaborative international effort designed to study basic hypersonic phenomena through flight experimentation. On 1 May 2012, the HIFiRE Flight 2 project successfully flew a hydrocarbon-fueled, Mach 8 scramjet experiment and demonstrated the ability to fly an accelerating, constant dynamic pressure trajectory consistent with a scramjet-powered flight vehicle using unguided sounding rocket techniques. The project goals of capturing high-quality flight data from a research scramjet operating through dual-to-scram mode transition up to and beyond Mach 8 were achieved. HIFiRE Flight 2 is unique in its contribution to scramjet research, providing a reference dataset to the hydrocarbon scramjet community while maturing a novel and reduced-cost strategy for performing similar future experiments. This paper describes the programmatic approach, the overall scramjet flight test experiment mission objectives, and the flight test strategy. It also includes an overview of launch system and payload hardware, and brief discussions of flight activities, flight data, and experimental results. Preliminary results from flight indicate a fully successful mission and experiment. Based on evaluations of data collected during the project, including ground test, flight test, and multiple analysis activities, all project-level scientific objectives have been successfully achieved.

Unsteady Three-Dimensional Mode Transition Phenomena in a Scramjet Flowpath

The physical process of mode transition, in which an accelerating hypersonic propulsion ramjet progresses from dual to scramjet mode operation, is explored in the context of the HIFiRE-2 ground and flight test campaign using three-dimensional, unsteady Reynolds-averaged Navier–Stokes simulations with a validated combustion model. Key highlights of the numerical procedure, mesh development, experimental validation, and the asymptotic end states are first summarized from prior work. From this basis, the methods and assumptions involved in the transient simulations are discussed with a focus on computational tractability. The results indicate that several major interactions play a crucial role in the mode transition, including separation associated with unsteady shock/turbulent boundary layer interactions, jet-in-crossflow barrel shock/flameholding, and cavity dynamics. The transition event, wherein the character of the flow changes in a rapid manner to delineate dual- and scramjet-mode phenomena, is generally similar in the flight and ground test simulations. As the freestream Mach number is increased under constant dynamic pressure and air-fuel ratio, shock interactions with the primary injectors result in loss of flameholding. The effects rapidly propagate downstream, altering the flow and combustion patterns in the cavity flameholder. Some noteworthy differences are present between the flight and ground test due to the shock system generated by the inlet of the flight article. These differences are highlighted to comment on the use of ground tests to predict flight performance. Details of the complex three-dimensional manner in which the combustion products and turbulence quantities vary during mode transition are also summarized.

Terminal Area Energy Management Trajectory Planning for an Unpowered Reusable Launch Vehicle with Gliding Limitations

RLVs' gliding capability, determined by its maximum dive and maximum range, provided a significant restriction in TAEM trajectory planning in this paper. The maximum-dive trajectory was generated based on Eq.(3) for a constant maximum dynamic pressure. In the guidance, it was optimized to be Eq.(13) for the open-loop command of bank angle in HAC segment. The simplified closed-loop command of angle of attack contained errors of altitude and path angle except the controlled velocity. Energy propagating as Eq.(8) calculated the reference velocity for the speed brake to track. Finally, an illustrative example was given to confirm the efficiency of the trajectory planning algorithm and optimized command. The simulation results in Fig.2 and Fig.3 indicate the proposed trajectory planning algorithm and guidance method are useful for the gliding capability limited RLV's TAEM with initial deviations.

Flexibility analysis of a tracking control method for air-breathing hypersonic vehicles

This paper studies the flexibility of a tracking control method originally proposed by the authors for air-breathing hypersonic vehicles (AHVs). The main feature of this method is to design the tracking controller without canceling but using aero-propulsive, as well as elevator-to-lift couplings. By introducing a virtual input, the tracking controller and external reference trajectories are simultaneously obtained by solving a system of linear algebraic equations. This system of linear algebraic equations is always solvable and the solution space of the corresponding homogeneous system is of dimension 3, which leads to much freedom in choosing or defining the free variables. The flexibility is reflected by the fact that the flight requirements of AHVs are involved in the definition of the free variables. Three case studies on different maneuvers, i.e., flight at constant dynamic pressure, flight at variant dynamic pressure and flight with fast climb rate are considered to verify the flexibility of this method. Simulation results show its effectiveness and flexibility.

Mission-Integrated Exergy Analysis for Hypersonic Vehicles: Methodology and Application

Recently developed theoretical work in which energy (the first law of thermodynamics) and entropy (the second law of thermodynamics) considerations are consistently applied to aerospace vehicles is used to provide a detailed exergy (availability) and performance analysis for an airbreathing hypersonic vehicle. An acceleration and climb mission at constant freestream dynamic pressure is performed with detailed instantaneous and time-integrated audits of entropy generation in and over the vehicle and in the vehicle wake. Entropy generation in the vehicle wake ranges from five to eight times the total entropy generation in and over the vehicle. The impact of irreversibility occurring in and over the vehicle itself on the total entropy generation in the wake is a small fraction of the overall wake losses. Fifteen percent of the overall energy input during the mission actually goes into productive acceleration and climb. The remainder is associated with the generation of entropy due to irreversibility in and over the vehicle and in the vehicle wake. The propulsion system is responsible for almost all entropy generation associated solely with the vehicle (excluding the overwhelmingly dominant contribution of the vehicle wake), and entropy generation in the combustor alone represents 85% of the total propulsion system loss.

Design of an Airbreathing Second Stage for a Rocket-Scramjet-Rocket Launch Vehicle

The most promising alternative to rockets for improved access to space involves staged systems using airbreathing propulsion. With scramjet technology improving, a number of airbreathing assisted access-to-space vehicle concepts have recently been proposed, including a three-stage rocket-scramjet-rocket launch architecture for payload masses on the order of 100 kg. This article presents a design methodology developed for the airbreathing second stage of such a system. This methodology uses multidisciplinary design optimization with simplified methods for the calculation of vehicle aerodynamics, propulsion, and mass. It has been applied to the design of a reusable scramjet-powered winged cone vehicle with a near-term Mach 6–12 hydrogen-fueled scramjet for propulsion. Through the manipulation of five vehicle design parameters, including the size and position of the engines, and flying the vehicle along constant dynamic pressure trajectories, a configuration was developed to maximize payload mass fraction to low Earth orbit. The mass of the whole system was 11.3 t, and that of the airbreathing vehicle was 4625 kg, delivering a 211 kg payload to low Earth orbit. This corresponds to an overall payload mass fraction of 1.87%. This methodology supplies a useful framework for developing a better understanding of the key drivers for airbreathing hypersonic accelerators.

Magnetopause expansions for quasi‐radial interplanetary magnetic field: THEMIS and Geotail observations

We report Time History of Events and Macroscale Interactions during Substorms (THEMIS) and Geotail observations of prolonged magnetopause (MP) expansions during long‐lasting intervals of quasi‐radial interplanetary magnetic field (IMF) and nearly constant solar wind dynamic pressure. The expansions were global: The magnetopause was located more than 3 RE and ∼7 RE outside its nominal dayside and magnetotail locations, respectively. The expanded states persisted several hours, just as long as the quasi‐radial IMF conditions, indicating steady state situations. For an observed solar wind pressure of ∼1.1–1.3 nPa, the new equilibrium subsolar MP position lay at ∼14.5 RE, far beyond its expected location. The equilibrium position was affected by geomagnetic activity. The magnetopause expansions result from significant decreases in the total pressure of the high‐β magnetosheath, which we term the low‐pressure magnetosheath (LPM) mode. A prominent LPM mode was observed for upstream conditions characterized by IMF cone angles less than 20°–25°, high Mach numbers and proton plasma β ≤ 1.3. The minimum value for the total pressure observed by THEMIS in the magnetosheath adjacent to the magnetopause was 0.16 nPa and the fraction of the solar wind pressure applied to the magnetopause was therefore 0.2, extremely small. The equilibrium location of the magnetopause was modulated by a nearly continuous wavy motion over a wide range of time and space scales.

On the Optimization of Aerospace Plane Ascent Trajectory

A hybrid heuristic optimization technique based on genetic algorithms and particle swarm optimization has been developed and tested for trajectory optimization problems with multi-constraints and a multi-objective cost function. The technique is used to calculate control settings for two types for ascending trajectories (constant dynamic pressure and minimum-fuel-minimum-heat) for a two-dimensional model of an aerospace plane. A thorough statistical analysis is done on the hybrid technique to make comparisons with both basic genetic algorithms and particle swarm optimization techniques with respect to convergence and execution time. Genetic algorithm optimization showed better execution time performance while particle swarm optimization showed better convergence performance. The hybrid optimization technique, benefiting from both techniques, showed superior robust performance compromising convergence trends and execution time.

Silanes as Fuels for Scramjets and Their Applications

In the light of recently revived interest in scramjet propulsion, a new look is being given at unconventional new fuels. Among the latter are, or could be included, hydrides. These contain hydrogen in variable amounts, suggesting their use for airbreathing applications in which a much higher density compared to that of L H 2 would be beneficial. In particular, this paper deals with silicon hydrides (silanes), because of their interesting combustion and energetic properties as fuels for scramjets. Silanes in combination with air seem, at this preliminary stage of analysis, an interesting conceptual alternative to L H 2 and, perhaps, also to LCH 4 for scramjet application. Accordingly, this paper explores the practical limits of application of silanes and their performance along cruiser and accelerator trajectories. Equilibrium composition of the combustion products of silanes, from monosilane up to pentasilane, were calculated. Ideal scramjet performance (specific thrust and specific impulse) were evaluated along a constant dynamic pressure trajectory from 21,350 to 30,500 m and compared to that with CH 4 /air and H 2 /air mixtures. High specific thrust is obtained as the equivalence ratio is increased; the I sp trend is the reverse, but still very appealing when weighted with the bulk density of silanes.