Freestream Dynamic Pressure Research Articles

Simplified Real-Time Flush Air-Data Sensing System for Sharp-Nosed Hypersonic Vehicles

This paper discusses a flush air-data sensing (FADS) system for a sharp-nosed hypersonic vehicle designed to estimate flight air data in real time at hypersonic speeds. The design’s target condition is Mach 5.0 to 7.0 with an angle of attack within 5 deg. The FADS system estimates air data by relating them to surface pressures measured from surface-mounted ports on the vehicle forebody. Unique combinations of pressure ratios were used to estimate the freestream dynamic pressure, which was a primary parameter for the flight experiment, the angle of attack, and the Mach number. The proposed FADS estimation algorithm was validated through numerical simulation, which was also used to generate datasets of surface pressures for given flight conditions. To handle possible sensor errors related to estimation accuracy in real-time estimation for a flight experiment, redundant systems were implemented. The results indicated that the designed FADS system can estimate air data within the uncertainty of 4.8% for a single estimator by considering sensor errors for freestream dynamic pressures in the range of 0–100 kPa, including targeted and off-design flight conditions. The proposed algorithm can estimate the air data with an acceptable level of uncertainty while retaining the robustness of estimation to sensor failures with low computational cost.

Coupling analysis of thin cropped delta wing at transonic speed based on partitioned approach

The oscillation of a thin cropped delta wing at transonic speed was investigated by fluid-structure interaction (FSI) analysis based on a coupling approach. The present analysis model was composed of open-source software fluid solver SU2, structure solver CalculiX, and coupling library preCICE. It was found that self-induced oscillation of the cropped delta wing with elastic deformation occurs at the transonic speed. The self-induced oscillation grew with the time elapsed, and then, the behavior reached limit-cycle oscillation (LCO). The times which reach LCO varied according to the freestream dynamic pressure. The primary frequency of the oscillation was approximately 55 Hz for all conditions considered herein. However, the eigenfrequencies of the present condition of the delta wing, which are 26.66, 89.12, 133.23, 204.60, and 312.34 Hz, etc., did not correspond to the oscillation frequency. It was indicated that the strong aerodynamic force induced the oscillation. Through the comparison with the experimental data, it was found that improvement of the present FSI analysis model, i.e., turbulence model and finer computational mesh, is required.

Pinhole measurements of hydroacoustic pressure fluctuations beneath wall-bounded turbulent flows

The common flush-mounted-transducer measurement method for underwater surface pressure fluctuations results in transducer-size-limited spatial resolution, which may be inadequate for the fluctuations produced beneath a turbulent boundary layer (TBL). For water moving over a smooth surface at speeds of a few meters per second, the relevant length scales may be less than 100 mm with fluctuation magnitudes far less than 1% of the free-stream dynamic pressure. Unfortunately, such size and sensitivity requirements cannot simultaneously be met with common instrumentation. However, mounting microphones behind pinholes is a possible solution for wall-bounded airflows. This presentation covers the effect of pinhole mounting pressure transducers for measuring surface pressure fluctuations in wall-bounded turbulent flows in water. The experimental work is conducted in a turbulent channel flow at varying flow speeds (0.5 to 7 m/s) using pressure transducers with a 5-mm-diameter sensitive area that are flush mounted and mounted behind variable size pinhole openings. The measured pressure-fluctuation power spectra are compared to each other and to results from past direct numerical simulations, in the 10 Hz to 10 kHz frequency range, to determine the utility of pinhole mounting for underwater TBL pressure-fluctuation measurements. [Sponsored by ONR.]

Experimental and Numerical Studies on Static Aeroelastic Behaviours of a Forward-Swept Wing Model

The static aeroelastic behaviours of a flat-plate forward-swept wing model in the vicinity of static divergence are investigated by numerical simulations and wind tunnel tests. A medium fidelity model based on the vortex lattice method (VLM) and nonlinear structural analysis is proposed to calculate the displacements of the wing structure with large deformation. Follower forces effect and geometric nonlinearity are considered to calculate the deformation of the wing by finite element method (FEM). In the wind tunnel tests, the divergence dynamic pressure is predicted by the Southwell method, and the static aeroelastic displacement is measured by a photogrammetric method. The results obtained by the medium fidelity model calculations show reasonable agreement with wind tunnel test results. A high fidelity model based on coupled computational fluid dynamics (CFD) and computational structural dynamics (CSD) predicts better results of the wing tip displacement when the freestream dynamic pressure is approaching the divergence dynamic pressure.

Flush Air Data Sensing System for a Sharp-Nosed Hypersonic Vehicle with Curved-Wedge Forebody

A new algorithm for a flush air data sensing (FADS) system was developed to estimate flight data and attitude angles for a sharp-nosed hypersonic vehicle in a framework of a hypersonic flight demonstration project directed by the Japan Aerospace Exploration Agency at the Kakuda Space Center. The design’s target condition is Mach 5.0–7.0 with angles of attack and sideslip within 6°. The FADS system relates air data and pressures from surface-mounted pressure ports on the forebody of the vehicle to generate lookup tables that are used to estimate the angle of attack, freestream Mach number, static pressure, and dynamic pressure. These data are also used to generate an angle-of-sideslip estimate from triples algorithm, which has been well-established for a blunt-nosed body. The estimation algorithm used by the designed FADS system is validated through numerical simulation, which was also used to generate lookup tables and wind tunnel experiments. Results indicated that the designed FADS system can estimate air data within the uncertainty of 6% for freestream conditions and within 0.7° for the angles of attack and sideslip in the range of targeted flight conditions.

Isolator characteristics under steady and oscillatory back pressures

Experiments are carried out to characterize the un-start/re-start phenomena in an isolator with steady and low-frequency oscillatory back pressures (0.75 Hz–2.5 Hz) in Mach 1.7 flow. In the present study, the effect of shock train interaction with the intake shocks is focused to capture these phenomena. It is demonstrated that the un-start mechanism is a continuous process with steady back pressures, whereas it is a discontinuous process under oscillatory back pressures. The un-start mechanism is triggered once the oscillatory back pressure is above the maximum isolator pressure. As the frequency of back pressure oscillation is increased, the isolator experiences an early un-start and a delayed re-start. On the other hand, if the oscillatory back pressure is decreased, the re-start process is initiated. During the re-start process, the time lag is found to be increased, and the hysteresis loss is decreased with an increase in oscillatory frequencies. These phenomena are studied through both the steady and unsteady pressure measurements along with instantaneous schlieren images for different dynamic pressures. Furthermore, the un-start can be avoided or delayed by increasing the freestream dynamic pressure.

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.

Unsteady pressure fluctuations measured on a hammerhead space vehicle and comparison with Coe and Nute's 1962 data

A detailed survey of pressure fluctuations was conducted on a replica of the generic, “hammerhead” shaped space vehicle tested more than half century ago by Coe and Nute (1962). A large number of current and past space vehicles follow this general hammerhead configuration. The present test was conducted in the 11-by-11 ft Transonic Wind Tunnel of NASA Ames Research Center over the critical transonic Mach number range of 0.6 ≤ M ≤ 1.2 where typical vehicles encounter very high levels of surface pressure fluctuations during ascent through the atmosphere. Out of the many different types of instrumentations used for this test, data from the dynamic pressure sensors are presented in this paper. In addition, shadowgraph flow visualizations, and a qualitative measure of turbulent fluctuations, estimated from high-speed photography, are also presented. The data set provided a clear insight into the local high level of fluctuations from the formation of the transonic shocks on the payload fairing and also from the impingement of the separated shear layer on the second stage. Space–time correlation showed interesting physics of upstream propagation of pressure fluctuations on the payload fairing when local shock waves were present. A comparison of the magnitude of pressure fluctuations between the present test and that done earlier by Coe and Nute showed good agreement when the limited frequency bandwidth of equipment used in the latter was accounted for. Similarly, when accounting for the differences in the free-stream dynamic pressure, the spectra of surface pressure fluctuations were in reasonable agreement between the present and Coe–Nute tests. The present test provides a far more complete picture of pressure fluctuations due to the usage of a large number of pressure transducers (216 versus 30), wider frequency bandwidth (50 kHz versus 1 kHz), and unsteady pressure-sensitive paint.

Effect of back pressure and freestream dynamic pressure on a typical Ramjet engine duct under realistic supersonic inlet condition

ABSTRACTThe present study investigates the behaviour of the shock train in a typical Ramjet engine under the influence of shock and expansion waves at the entry of a low aspect ratio (1:0.75) rectangular duct/isolator at supersonic Mach number (M = 1.7). The start/unstart characteristics are investigated through steady/unsteady pressure measurements under different back and dynamic pressures while the shock train dynamics are captured through instantaneous Schlieren flow visualisation. Two parameters, namely pressure recovery and the pressure gradient, is derived to assess the duct/isolator performance. For a given back pressure, with maximum blockage (9% above nominal), the duct/isolator flow is established when the dynamic pressure is increased by 23.5%. The unsteady pressure measurements indicate different scales of eddies above 80 Hz (with and without flap deflection). Under the no flap deflection (no back pressure) condition, the maximum fluctuating pressure component is 0.01% and 0.1% of the stagnation pressure at X/L = 0.03 (close to the entry of the duct) and X/L = 0.53 (middle of the duct), respectively. Once the flap is deflected (δ = 8°), decay in eddies by one order is noticed. Further increase in back pressure (δ ≥ 11°) leads the flow to unstart where eddies are observed to be disappeared.

Conceptual Modeling of Drag-Augmented Supersonic Retropropulsion for Mars Entry, Descent, and Landing

The development of new decelerator technologies will be required as the payload mass for future Mars landing missions increases beyond the current state-of-the-art capability. This study examines the potential for supersonic retropropulsion applied on entry, descent, and landing vehicles to increase the landed payload mass. This study describes the development of a model characterizing the drag augmentation capabilities of peripheral-nozzle supersonic retropropulsion flow interactions. The model captures the dominant flow physics of pressure conservation through shock cascade structures and predicts an increase in the drag coefficient over the nominal drag coefficient of a 70 deg sphere-cone aeroshell by 14% at high Mach numbers. This study also describes drag-augmented supersonic retropropulsion operation concepts for use in Mars entry, descent, and landing. Drag-augmented supersonic retropropulsion is found to be most effective when used in the region of maximum freestream dynamic pressure. The vehicle dry mass is increased by 47% over the reference ballistic trajectory. The region of influence for aerodynamic–propulsive interactions is identified for a set of constant-thrust supersonic retropropulsion trajectories. A hybrid concept combining supersonic retropropulsion and an inflatable aerodynamic decelerator is found to be capable of providing vehicle dry masses that are 707% larger than the baseline vehicle studied.

Downstream flow structures of a fluttering piezoelectric energy harvester

The flow structures downstream of a fluttering piezoelectric energy harvester are examined, since 40% increases in power output have been noted from a downstream harvester. Smoke wire flow visualisation with a novel bi-lighting technique, and simultaneous voltage measurements are used to understand this phenomenon. The bi-lighting technique helped to confirm the out-of-plane velocity fields caused by the tilted cone vortices shed from the upstream harvester. It was speculated that the time-varying local pressure gradients induced by the propagating vortices coupled beneficially with the free-stream dynamic pressure and flutter instabilities, acting to elevate the flutter amplitude of the downstream harvester. Also, the phase lag between the harvesters showed a nonlinear relationship with the free-stream wind speed.

FADS系统在尖楔前体高超声速飞行器中的应用

For some hypersonic flight vehicles with sharp wedged fore-bodies, especially for those which are used to verify supersonic combustion ramjet technology, the precision of angle of attack will not be achieved only through Inertial Navigation System (INS). The aerodynamic model for Flush Air Data Sensing System (FADS) is built based on the surface pressure distribution obtained through the pressure orifices laid on specific positions of the surface, and the flight parameters, such as angle of attack, angle of side-slip, Mach number, free-stream static pressure and dynamic pressure are inferred from the aerodynamic model. For FADS/INS integrated system, angle of attack can be determined through aerodynamic model we have firstly built based on oblique shock wave theory, and the error for angle of attack is accurate enough (<0.1°). For FADS system alone, the flight parameters can be determined through aerodynamic model based on two and three dimensional flow theory, and the error for Mach number is about 5%. The results show that FADS system can be applied widely in hypersonic flight vehicles with sharp wedged fore-bodies.

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.

Trajectory Estimation of the Hayabusa Spacecraft During Atmospheric Disintegration

The Hayabusa spacecraft was intentionally destroyed in the atmosphere at superorbital velocity at the conclusion of its asteroid sample return mission in June 2010. This study uses single-station ground-based video observations of the reentry to analyze the breakup of the spacecraft and estimate the trajectory of 80 individual spacecraft fragments. An extended Kalman filter with batch initialization is used to estimate the position, velocity, and aerodynamic ballistic coefficients of the fragments. The breakup is characterized and compared with preflight predictions. A high area-to-mass object is seen early during the reentry, which matches closely with the predicted solar panel separation. Nearly all fragments have decreasing freestream dynamic pressure during their observed trajectories. Fragments with high drag ballistic coefficients are more likely to be observed early in the reentry. Assuming simple aluminum spheres, the estimated ballistic coefficients show that the fragments would have radius and mass on the order of centimeters and grams, respectively. The ablation coefficient is also calculated for several fragments using the estimated velocities and ballistic coefficients; results resemble natural meteors and agree with other studies of the Hayabusa reentry using different methods.

Effects of streamwise rotation on the dynamics of a droplet

An initially streamwise rotating droplet released into a uniform cross flow is studied numerically. The computations are performed using a finite volume Navier–Stokes solver which employs a moving mesh interface tracking scheme to locate the interface. With a large initial Weber number (Wei = 40) the streamwise rotating droplet flattens along the free stream direction more quickly as rotation rate (\documentclass[12pt]{minimal}\begin{document}$\varOmega ^*$\end{document}Ω*) increases, and leads to a dramatic increase in the dynamic drag coefficient (CD/A*, where A* is the dimensionless frontal area). On the other hand, for Wei = 4 and 0.4 at \documentclass[12pt]{minimal}\begin{document}$\varOmega ^* \ge 0.6$\end{document}Ω*≥0.6, the flattening of the droplet is less pronounced and the droplet even restores to spherical shape, hence, CD/A* decreases sharply. The dynamic drag coefficient even becomes negative for Wei = 4 and 0.4 at \documentclass[12pt]{minimal}\begin{document}$\varOmega ^* = 1$\end{document}Ω*=1. At the largest deformation, the droplet can be classified into three major shapes: biconvex, convex-concave, and biconcave. One dominant feature of the wake downstream of the droplet is the formation and convection of vortex rings. The shape and deformation of the droplet is dependent not only on the size of the vortex ring, but also upon the free stream dynamic pressure and droplet pressure. The detachment of vortex ring in the wake leads to a substantial drag reduction, and this detachment occurs at Re ≈ 28.

Prediction of total pressure characteristics in the settling chamber of a supersonic blowdown wind tunnel

AbstractOccurrence of transient starting and stopping loads during tests at high Mach numbers is one of the major problems in intermittent blowdown wind tunnels. It is believed that in order to overcome this problem, the wind tunnel could be started at a low Mach number and low stagnation pressure; the desired high Mach number condition could be reached by continuously changing the nozzle contour while synchronously increasing the stagnation pressure. After completing the tests, the nozzle could be brought back to the initial low Mach number accompanied by synchronous decrease in the stagnation pressure. In such a scenario, it is important to ensure that the pressure regulating valve (PRV) of the wind tunnel delivers and maintains a specified minimum stagnation pressure at any Mach number, so that supersonic breakdown of the test section flow does not occur. In this paper, the problem is formulated based on quasi-steady one-dimensional isentropic equations and numerically solved to predict the time histories of settling chamber pressure and storage tank pressure for a given trajectory of the opening of the PRV, as the Mach number is changed from Mach 1 to 4·0 continuously in four seconds and vice versa. The effects of rate of change of PRV open area and rate of change of Mach number on the stagnation pressure characteristics in the settling chamber and storage tank are predicted. The measured trajectories of the PRV in experiments in the NAL 0·6m transonic wind tunnel are used as input to the prediction program to validate the methodology. Predictions indicate that when the nozzle throat is changed from Mach 1 to 4 in four seconds, the settling chamber stagnation pressure rapidly builds up and approaches the pressure in the storage tank. Predictions show an alarming rise in free stream dynamic pressure during transition from Mach 1 to 4 and vice versa, which needs to be verified through measurements.

The wall pressure signature of transonic shock/boundary layer interaction

The structure of wall pressure fluctuations beneath a turbulent boundary layer interacting with a normal shock wave at Mach number M∞ = 1.3 is studied exploiting a direct numerical simulation database. Upstream of the interaction, in the zero-pressure-gradient region, pressure statistics compare well with canonical low-speed boundary layers in terms of fluctuation intensities, space–time correlations, convection velocities and frequency spectra. Across the interaction zone, the root-mean-square wall pressure fluctuations attain very large values (in excess of 162 dB), with a maximum increase of about 7 dB from the upstream level. The two-point wall pressure correlations become more elongated in the spanwise direction, indicating an increase of the pressure-integral length scales, and the convection velocities (determined from space–time correlations) are reduced. The interaction qualitatively modifies the shape of the frequency spectra, causing enhancement of the low-frequency Fourier modes and inhibition of the higher ones. In the recovery region past the interaction, the pressure spectra collapse very accurately when scaled with either the free-stream dynamic pressure or the maximum Reynolds shear stress, and exhibit distinct power-law regions with exponent −7/3 at intermediate frequencies and −5 at high frequencies. An analysis of the pressure sources in the Lighthill's equation for the instantaneous pressure has been performed to understand their contributions to the wall pressure signature. Upstream of the interaction the sources are mainly located in the proximity of the wall, whereas past the shock, important contributions to low-frequency pressure fluctuations are associated with long-lived eddies developing far from the wall.

Optimal Movable Wing Tip Strake

Nomenclature AR = aspect ratio, b=S b = span, mm CD = configuration drag coefficient CL = configuration lift coefficient c = chord, mm ds = angle of strake setting relative to wing chord plane (positive, strake leading edge up), q = freestream dynamic pressure, 1 2 1V 2 1, kPa Re = Reynolds number based on wing model chord, 1V1c= 1 S = reference area, m V1 = freestream velocity, m=s x, y, z = aerodynamic axes = angle of attack, deg = absolute viscosity, N s=m = air density, kg=m

New visions in experimental investigations of a supersonic under-expanded jet into a high subsonic crossflow

A series of experiments was performed to investigate the interaction of an under-expanded axisymmetric supersonic jet exhausted from a flat plate with a high subsonic crossflow. The goal was to study the effect of boundary layer thickness (δ) and jet to freestream dynamic pressure ratio ( J) on flow field pressure distributions. The resulting measurements upstream of the jet showed that with increasing boundary layer thickness, the magnitude of the pressure coefficient decreases, whereas downstream of the jet, the recovery of the back-pressure moved closer to the nozzle exit. Flow field measurements indicated that with increasing boundary layer thickness, the jet plume dissipation rate increased, whereas the strength of the counter-rotating vortex pair (CRVP) did not vary significantly. In addition, it was clearly observed that with increasing J, the CRVP penetrated into the crossflow and the magnitude of the pressure coefficient on the surface upstream and downstream of the jet increased.

Examining Effects of Increased Effective Area on Performance of Movable Tip Strakes

Nomenclature  = b=S, aspect ratio b = span, mm CD = wing drag coefficient CL = wing lift coefficient c = chord, mm d = angle relative to wing chord plane, deg q = 1=2 1V 1, freestream dynamic pressure, kPa Re = 1V1c= 1, Reynolds number based on wing model chord S = reference area, mm V1 = freestream velocity, m=s x, y, z = aerodynamic axes = angle of attack, deg = strake leading-edge sweep angle, deg = absolute viscosity, N s=m = air density, kg=m