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The reconstruction of atmospheric properties encountered during Mars entry trajectories is a crucial element of postflight mission analysis. This paper proposes a deep learning architecture using a long short-term memory (LSTM) network for the reconstruction of Martian density and wind profiles from inertial measurements and guidance commands. The LSTM is trained on a large set of Mars entry trajectories controlled through the fully numerical predictor-corrector entry guidance (FNPEG) algorithm, with density and wind from the Mars Global Reference Atmospheric Model (GRAM) 2010. The training of the network is examined, ensuring that the LSTM generalizes well to samples not present in the training set, and the performance of the network is assessed on a separate training set. The errors of the reconstructed density and wind profiles are, respectively, within 0.54 and 1.9%. Larger wind errors take place at high altitudes due to the decreased sensitivity of the trajectory in regions of low dynamic pressure. The LSTM architecture reliably reproduces the atmospheric density and wind encountered during descent.

Experiments with a scramjet engine for a planned Mach 5–10 access-to-space vehicle have been undertaken in the T4 reflected shock tunnel at low Mach number conditions. The 3D scramjet flowpath is designed to integrate with a conical vehicle and involves significant shape transition to enhance aerodynamic efficiency. Gaseous hydrogen is used as the fuel, injected at multiple locations through the engine. A generic thrust nozzle is included downstream of the diverging section of the combustor. Distributed injection is shown to be effective at a Mach 7 flight condition with a nominal dynamic pressure of 55 kPa. For an increased dynamic pressure of 130 kPa, robust separated combustion is observed at lower total equivalence ratios and without the need for distributed injection. At a simulated Mach 6 flight condition, no ignition is observed at 55 kPa while robust combustion occurs at the high pressure condition. These experiments utilised a flowpath which was altered from an operational flight design to fit within the tunnel core flow. Hence, the experiments are exploratory in nature rather than demonstrating the performance of an operational engine.

X3R is a new, large-scale free-piston driven reflected shock tunnel designed and operated at the University of Queensland, Australia. The facility is a derivative of the X3 super-orbital expansion tube, which uses a combination of new and pre-existing hardware. It is capable of generating 600 mm diameter Mach 7 test flows at 25 kPa and 50 kPa dynamic pressure for durations of approximately 12.5 ms. Previously, hypersonic testing in Australia was limited in both scale and duration, due to the size of the local facilities capable of experimentation in this regime. This paper presents the new facility and summarizes its current operational capabilities. This includes a description of the facility geometry, and details of the two currently available operating conditions. These conditions were specifically designed to overcome the drawbacks of the tunnel’s relatively short free-piston driver, and use a high helium fraction driver gas to produce a long duration but low compression ratio driver response.

The qualification of the Orion Capsule Parachute Assembly System (CPAS) system includes exposure to loads and dynamic pressures above the required values as validation of the parachutes' structural integrity. As outlined in the certification plan, each of the four parachutes of the system are to be subjected to 110% of their respective maximum dynamic pressure requirements. The Main and Drogue parachutes have satisfied this overload condition in drop testing and due to cost and schedule constraints, the Forward Bay Cover Parachute (FBCP) and Pilot parachute were subjected to the overload condition in the ground testing described in this document. The test objectives and pass/fail criteria were established and require the parachutes to achieve and maintain a target riser load (associated with a minimum of 110% dynamic pressure overload) for a minimum of three seconds while sustaining no failures of any structural members (vent hoop, radials, suspension line) or any damage which propagates into catastrophic failure of the canopy. Considering the assumptions and limitations associated with the ground testing (primarily non-uniform flow field of the ground test system and variations in parachute manufacturing), a method of establishing the desired overload condition was determined by the technical community and covers the 2-sigma bounds of the drag area distribution derived from drop testing. On June 27, 2017 the testing was executed at the High Velocity Airflow System (HiVAS) facility located at the Weapon Survivability Laboratory (WSL) at the Naval Air Warfare Center, Weapons Division (NAWCWD) China Lake. Engineering Development Unit (EDU) parachutes were used as pathfinders to gain experience with achieving the test conditions. Additional runs were executed to measure the airflow at the same location as the canopy skirt, although this data is not required to satisfy the test objectives and parachute pass/fail criteria. The qualification parachutes were successfully exposed to the target conditions and sustained only minor damage.

In an extensive measurement campaign at the UFFA test rig of AneCom AeroTest in Wildau, Germany, the sound field generated from the ACAT1 fan stage was measured by several dynamic pressure sensor arrays of different geometrical arrangements at the inlet, interstage and bypass section. The sparse ring arrays with irregular and optimised geometries are utilised in combination with the Compressed Sensing-based mode analysis in order to investigate the tonal and broadband sound generation and transmission with high resolution. From the azimuthal mode spectra, the relation between the cut-on mode range at the sources and the dominant modes at the measurement sections is observed and used to determine the contribution of different sound generation mechanisms. With the help of interstage mode analyses strong rotor-coherent contributions due to rotor wake irregularities are identified and the rotor-shielding effect is assessed. The dependencies of the broadband mode spectra with respect to the operation points and working lines are shown and the contribution of stator and rotor bound sources is investigated. Results indicate that mode scattering at the stator vane row has a significant influence on the downstream radiated sound field. Evidence is further found for the occurrence of the acoustic tunnel effect, where cut-off modes excited at the stator leading edge become cut-on during the propagation through the inter-vane channels. This effect triggers mode scattering towards rotor counter-rotating modes and increases the sound radiation from fan stages.

Flush air data sensing (FADS) systems have been previously used at both Earth and Mars to provide onboard estimates of angle of attack, sideslip angle, and dynamic pressure. However, these FADS data were often not used in an in-the-loop sense to inform the onboard guidance and control systems. A method to integrate FADS-derived density and wind estimates with a numerical predictor-corrector guidance algorithm is presented. The method is demonstrated in a high-fidelity simulation of a human-scale Mars entry vehicle that utilizes a hypersonic inflatable aerodynamic decelerator (HIAD) with direct force control. Effects on guidance commands and state uncertainties both with and without FADS system modeling are presented and discussed.

Improvements and results of a new method are presented that computes a pre-test estimate of the precision error of the drag coefficient of a wind tunnel model. The error estimate is defined as the part of the drag coefficient's precision error that is primarily associated with the precision error of the angle of attack measurement and physical characteristics of the chosen strain-gage balance. The method indirectly describes the precision error of the angle of attack measurement by using an assumed balance gage output variation of one microV/V. The physical characteristics of the balance, on the other hand, are described by partial derivatives of the axial and normal forces with respect to the strain-gage outputs. These derivatives can directly be obtained from the data reduction matrix of the balance. The precision error estimate itself is calculated by applying a simple explicit equation that uses the model reference area, the dynamic pressure, the angle of attack, the coefficients of the linear terms of the data reduction matrix, and the electrical output variation of one microvolt per volt as input. Precision errors at constant angle of attack may be visualized as contour plots by plotting them, for example, versus the Mach number and the total pressure. Characteristics of NASA's MC60E balance are used in combination with the reference area of a generic wind tunnel model in order to demonstrate that error estimates are independent of both the balance load format and the units chosen for the description of balance loads, model reference area, and the dynamic pressure. Finally, experimental data from a wind tunnel test of the Ames Check Standard Model in the NASA Ames 11-foot Transonic Wind Tunnel illustrates the application of the method to real-world test data.

The Advanced Supersonic Parachute Inflation Research and Experiments project is a flight test program for development of supersonic parachutes for future use at Mars. The flight tests are a risk-reduction program for the Mars 2020 mission. The flight tests involve two Disk-Gap- Band parachute designs to be tested at relevant Mach number and dynamic pressure regimes for the Mars 2020 entry capsule. The first of these parachutes is a built-to-print design that was successfully employed by the Mars Science Laboratory lander at Mars in August 2012, and the second is a design that is strengthened in material properties and construction methods but has the same geometry as that used by Mars Science Laboratory. The first flight test of the built-to-print parachute took place on October 4, 2017 at NASA's Wallops Flight Facility. The parachute test was successful. The second and third flight tests took place on March 31st and September 7th 2018, respectively, and successfully tested the strengthened parachute design. This paper describes the instrumentation, data analysis techniques, and atmospheric and trajectory reconstruction results from the second and third flight tests.

An experimental investigation of the combustion dynamic characteristics of two advanced multi-cup lean direct injectors (LDI) under simulated gas turbine combustor conditions was conducted. The objective was to gain a better understanding of the physical phenomena inside a pressurized flame tube combustion chamber and study the effects of injector flow number on combustion dynamics. The injectors are known as Three-zone Injectors one and two or 3ZI-1 and 3ZI-2, respectively. The injectors were experimentally evaluated at inlet pressures up to 1.724 MPa, non-vitiated air temperatures up to 828K, and adiabatic flame temperatures up to 1975K. Dynamic pressure measurements were taken upstream of the injectors and in the combustion zone. The combustion dynamic behavior of the two injectors was measured over a range of inlet pressures, inlet temperatures, fuel air ratios, and fuel flow splits.

The Advanced Supersonic Parachute Inflation Research and Experiments project is a flight test program for development of supersonic parachutes for potential future use at Mars. The flight tests are designed to reduce risk for the Mars 2020 mission. The flight tests involve two Disk-Gap-Band parachute designs to be tested at relevant Mach number and dynamic pressure conditions for the Mars 2020 entry capsule. The first of these parachutes is a built-to-print design that was successfully employed by the Mars Science Laboratory lander at Mars in August 2012, and the second is a design that is strengthened in material properties and construction methods but has the same geometry as that used by Mars Science Laboratory. The first flight test of the built-to-print parachute took place on October 4, 2017 at NASA's Wallops Flight Facility. This paper describes the instrumentation, data analysis techniques, and atmospheric and trajectory reconstruction results from this flight test.

The NASA Learn-to-Fly (L2F) project recently completed a series of flight demonstrations of its learning algorithm for flight control at Fort A. P. Hill in Virginia. This paper discusses the test setup and concept of operations (ConOps) used by the L2F team. Unmanned airframe demonstrators for testing the research algorithms included a modified commercial off-the-shelf subscale powered airplane, plus four gliders – two of which had an unconventional configuration and were fabricated using a “rapid” prototyping technique. Avionics system similarities and differences between the test aircraft are described, as well as ground testing in preparation for flight. The ConOps discussion includes the development of a tethered helium balloon drop launch technique for the glider demonstrators. This launch method was chosen for its potential to be inexpensive and allow for rapid turn-around for multiple glider launches – but it also presented challenges, such as balloon tether avoidance, high angle of attack, low dynamic pressure initial conditions, and susceptibility to winds. A remotely piloted approach employing high-end hobbyist radio controlled (R/C) hardware was used for the powered demonstrator. This approach accommodated the interaction between the R/C flight system and the research flight control computer, engaging the L2F algorithm at varying initial conditions and artificially reducing the aircraft stability to stress the algorithm.

This paper describes a lightweight, large-area solar array concept for Mars surface power called the Compact Telescoping Surface Array (CTSA). The design is derived from the Compact Telescoping Array (CTA) proposed in 2015 for high-power spacecraft. The CTSA deploys horizontally from Mars landers, provides 1000 sq.m of solar cell area, and generates about 50-80 kW daytime and 10 kW nighttime power (from energy storage) near the equator with clear skies. The total mass is about 1500 kg, and the stowed volume is about 10 cu.m, equivalent to 200 W/kg and 30 kW/cu.m at 1 astronomical unit (AU) from the sun. These challenging mass and volume goals require compact, lightweight structures to deploy and tension thin solar cell blankets in the Mars 0.38 g gravity field. Other features include single-axis rotation for daily or seasonal sun tracking, dust shedding, and feathering in high winds and sufficient strength for sustained 40 m/s winds in dust storms and brief 100 m/s winds in dust devils (dynamic pressure of 120 Pa). Based on preliminary design and analysis reported in the paper, CTSA appears to be a feasible approach for 1000 sq.m-class Mars solar arrays that autonomously deploy from landers prior to human arrival.

© 2019 by Max Opgenoord, Mark Drela, and Karen Willcox. Transonic aeroelasticity is an important consideration in the conceptual design of next-generation aircraft configurations. This paper develops a low-order physics-based flutter model for swept high-aspect-ratio wings. The approach builds upon a previously developed flutter model that uses the flowfield's lowest moments of vorticity and volume-source density perturbations as its states. The contribution of this paper is a new formulation of the model for swept high-aspect-ratio wings. The aerodynamic model is calibrated using offline two-dimensional unsteady transonic computational-fluid-dynamics simulations. Combining that aerodynamic model with a beam model results in a low-dimensional overall aeroelastic system. The low computational cost of the model permits its incorporation in a conceptual design tool for next-generation transport aircraft. The model's capabilities are demonstrated by finding transonic flutter boundaries for different clamped-wing configurations and investigating the influence of transonic flutter on the planform design of next-generation transport aircraft.

The new check standard model of the NASA Ames 11-ft Transonic Wind Tunnel was chosen for a future validation of the facility's wall interference correction system. The chosen validation approach takes advantage of the fact that test conditions experienced by a large model in the slotted part of the tunnel's test section will change significantly if a subset of the slots is temporarily sealed. Therefore, the model's aerodynamic coefficients have to be recorded, corrected, and compared for two different test section configurations in order to perform the validation. Test section configurations with highly accurate Mach number and dynamic pressure calibrations were selected for the validation. First, the model is tested with all test section slots in open configuration while keeping the model's center of rotation on the tunnel centerline. In the next step, slots on the test section floor are sealed and the model is moved to a new center of rotation that is 33 inches below the tunnel centerline. Then, the original angle of attack sweeps are repeated. Afterwards, wall interference corrections are applied to both test data sets and response surface models of the resulting aerodynamic coefficients in interference-free flow are generated. Finally, the response surface models are used to predict the aerodynamic coefficients for a family of angles of attack while keeping dynamic pressure, Mach number, and Reynolds number constant. The validation is considered successful if the corrected aerodynamic coefficients obtained from the related response surface model pair show good agreement. Residual differences between the corrected coefficient sets will be analyzed as well because they are an indicator of the overall accuracy of the facility's wall interference correction process.

Space maneuver vehicles (SMVs) [1,2] will play an increasingly important role in the future exploration of space because their on-orbit maneuverability can greatly increase the operational flexibility, and they are more difficult as a target to be tracked and intercepted. Therefore, a well-designed trajectory, particularly in the skip entry phase, is a key for stable flight and for improved guidance control of the vehicle [3,4]. Trajectory design for space vehicles can be treated as an optimal control problem. Because of the highly nonlinear characteristics and strict path constraints of the problem, direct methods are usually applied to calculate the optimal trajectories, such as the direct multiple shooting method [5], direct collocation method [5,6], or hp hp -adaptive pseudospectral method [7,8].

This presentation is a conference companion and discusses a new flutter analysis technique using CFD results.

This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing (BSCW) configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results show the effects of the temporal and spatial resolution, the coupling scheme between the flow and the structural solvers, and the initial excitation conditions on the numerical flutter onset. Depending on the free stream condition and the angle of attack, the above parameters do affect the flutter onset. Two conditions are analyzed: Mach 0.74 with angle of attack 0 and Mach 0.85 with angle of attack 5. The results are presented in the form of the damping values computed from the wing pitch angle response as a function of the dynamic pressure or in the form of dynamic pressure as a function of the Mach number.

A new method is presented that computes a pre{test estimate of the upper bound of the drag coefficient repeatability of a wind tunnel model. This upper bound is a conservative estimate of the precision error of the drag coefficient. For clarity, precision error contributions associated with the measurement of the dynamic pressure are analyzed separately from those that are associated with the measurement of the aerodynamic loads. The upper bound is computed by using information about the model, the tunnel conditions, and the balance in combination with an estimate of the expected output variations as input. The model information consists of the reference area and an assumed angle of attack. The tunnel conditions are described by the Mach number and the total pressure or unit Reynolds number. The balance inputs are the partial derivatives of the axial and normal force with respect to all balance outputs. Finally, an empirical output variation of 1.0 microV/V is used to relate both random instrumentation and angle measurement errors to the precision error of the drag coefficient. Results of the analysis are reported by plotting the upper bound of the precision error versus the tunnel conditions. The analysis shows that the influence of the dynamic pressure measurement error on the precision error of the drag coefficient is often small when compared with the influence of errors that are associated with the load measurements. Consequently, the sensitivities of the axial and normal force gages of the balance have a significant influence on the overall magnitude of the drag coefficient's precision error. Therefore, results of the error analysis can be used for balance selection purposes as the drag prediction characteristics of balances of similar size and capacities can objectively be compared. Data from two wind tunnel models and three balances are used to illustrate the assessment of the precision error of the drag coefficient.

An aeroelastic analysis of a prototype distributed electric propulsion wing is presented. Results using MSC Nastran (Registered Trademark) doublet lattice aerodynamics are compared to those based on FUN3D Reynolds Averaged Navier- Stokes aerodynamics. Four levels of grid refinement were examined for the FUN3D solutions and solutions were seen to be well converged. It was found that no oscillatory instability existed, only that of divergence, which occurred in the first bending mode at a dynamic pressure of over three times the flutter clearance condition.

Combustion dynamics data were collected at the NASA Glenn Research Center's CE-5 flame tube test facility under combustor outlet choked conditions. Two 9-point Swirl-Venturi Lean Direct Injection (SV-LDI) configurations were tested in a rectangular cuboid combustor geometry. Combustion dynamic data were measured at different engine operational conditions up to inlet air pressure and temperature of 24.13 bar and 828 K, respectively. In this study, the effects of acoustic cavity resonance, precessing vortex core (PVC), and non-uniform thermal expansion on the dynamic noise spectrum are identified by comparing the dynamic data that collected at various combustor inlet conditions along with combustor geometric calculations. The results show that the acoustic cavity resonance noises were seen in the counter-rotating pilot configuration but not in the co-rotating pilot configuration. Dynamic pressure noise band at around 0.9 kHz was only detected at the P'41 location (9.8 cm after fuel injector face) but not at the P'42 location (29 cm after the fuel injector face); the amplitude of this noise band depended on the thermal expansion ratio (T4/T3). The noise band at around 1.8 kHz was found to depend on the inlet air pressure or the air density inside the combustor. The PVC frequency was not observed in these two configurations.

CFD analysis is presented of the mixing characteristics and performance of three fuel injectors at hypervelocity flow conditions. The calculations were carried out using the VULCAN-CFD solver and Reynolds-Averaged Simulations (RAS). The high Mach number flow conditions match those proposed for the planned experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships relevant to flight Mach numbers greater than eight. Because of the high Mach number flow considered, the injectors consist of a fuel placement device, a strut; and a fluidic vortical mixer, a ramp. These devices accomplish the necessary task of distributing and mixing fuel into the supersonic cross-flow albeit via different strategies. Both of these devices were previously studied at lower flight Mach numbers where they exhibited promising performance in terms of mixing efficiency and total pressure recovery. For comparison, a flush-wall injector is also included. This type of injector generally represents the simplest method of introducing fuel into a scramjet combustor, however, at high flight Mach number conditions, the dynamic pressure needed to induce sufficient fuel penetration may be difficult to achieve along with other requirements such as achieving desired levels of fuel-to-air mixing at the required equivalence ratio. The three injectors represent the baseline configurations planned for the experiments. The current work discusses the mixing flow field behavior and differences among the three fuel injectors, mixing performance as described by the mixing efficiency and the total pressure recovery, and performance considerations based on the thrust potential.

NASA’s Low-Density Supersonic Decelerators project (LDSD) has developed and tested four new aerodynamic decelerator technologies for future Mars missions: two attached toroidal inflatable decelerators, a ballute, and a large supersonic parachute. On June 8, 2015, the project conducted a high-altitude, supersonic flight test of a 30.5-meter supersonic Ringsail (SSRS) canopy at the US Navy’s Pacific Missile Range Facility (PMRF) on Kauai, HI. This test, the second in a series of Supersonic Flight Dynamics Tests (SFDT-2), allowed the LDSD project to test the deployment and performance of its parachute decelerator system in the wake of a representative test vehicle (a 4.7-meter aeroshell and 6-meter toroidal inflatable aerodynamic decelerator) at conditions relevant to Mars entry for the second time. The parachute decelerator system consisted of the SSRS main parachute and a 4.4-meter ballute (called the parachute deployment device, or PDD) for its extraction. The ballute was mortar-deployed at a Mach number of 2.78 and a dynamic pressure of 493 Pa, and inflated with the aid of a water-methanol based gas generator. After flying in the wake of the test vehicle for ten seconds, the PDD was released and allowed to extract the main parachute pack. The SSRS reached line-stretch at a Mach number of 2.37 and dynamic pressure of 602 Pa. Following full inflation, the propagation of a tear in the canopy led to the failure of the parachute skirt band and to the subsequent failure of the vent band. The test vehicle was instrumented with load sensors, inertial sensors, and high-speed and high resolution cameras that provided data on the performance of the PDD and SSRS through deployment, inflation, and flight. This paper describes the resulting reconstructed behavior of the PDD and SSRS during deployment and inflation, their aerodynamic performance on SFDT-2, the failure of the SSRS shortly after full inflation, and the LDSD project’s investigation into its underlying causes.

A detailed uncertainty analysis has been performed to determine the measurement uncertainty of the freestream parameters and the force and moment coefficients at the NRC 1.5-meter trisonic wind tunnel. The systematic and random components are investigated separately using the Taylor series propagation method. Detailed breakdowns of the uncertainties have been obtained in order to identify the variables that have the largest contributions to the uncertainty of the test results. A summary of the findings is presented for Mach number, dynamic pressure and coefficients of drag and lift, under different operating conditions., 32nd AIAA Aerodynamic Measurement Technology and Ground Testing Conference, June 13-17, 2016, Washington, D.C. USA

The recent advancement in fast-response Pressure-Sensitive Paint (PSP) allows time-resolved measurements of unsteady pressure fluctuations from a dense grid of spatial points on a wind tunnel model. This capability allows for direct calculations of the wavenumber-frequency (k-ω) spectrum of pressure fluctuations. Such data, useful for the vibro-acoustics analysis of aerospace vehicles, are difficult to obtain otherwise. For the present work, time histories of pressure fluctuations on a flat plate subjected to vortex shedding from a rectangular bluff-body were measured using PSP. The light intensity levels in the photographic images were then converted to instantaneous pressure histories by applying calibration constants, which were calculated from a few dynamic pressure sensors placed at selective points on the plate. Fourier transform of the time-histories from a large number of spatial points provided k-ω spectra for pressure fluctuations. The data provides first glimpse into the possibility of creating detailed forcing functions for vibro-acoustics analysis of aerospace vehicles, albeit for a limited frequency range.

The relationship between the aerodynamic lift force generated by a wind tunnel model, the model weight, and the measured normal force of a strain-gage balance is investigated to better understand the expected use envelope of the normal force during a wind tunnel test. First, the fundamental relationship between normal force, model weight, lift curve slope, model reference area, dynamic pressure, and angle of attack is derived. Then, based on this fundamental relationship, the use envelope of a balance is examined for four typical wind tunnel test cases. The first case looks at the use envelope of the normal force during the test of a light wind tunnel model at high subsonic Mach numbers. The second case examines the use envelope of the normal force during the test of a heavy wind tunnel model in an atmospheric low-speed facility. The third case reviews the use envelope of the normal force during the test of a floor-mounted semi-span model. The fourth case discusses the normal force characteristics during the test of a rotated full-span model. The wind tunnel model's lift-to-weight ratio is introduced as a new parameter that may be used for a quick pre-test assessment of the use envelope of the normal force of a balance. The parameter is derived as a function of the lift coefficient, the dimensionless dynamic pressure, and the dimensionless model weight. Lower and upper bounds of the use envelope of a balance are defined using the model's lift-to-weight ratio. Finally, data from a pressurized wind tunnel is used to illustrate both application and interpretation of the model's lift-to-weight ratio.

Concept, design and application of a quasi-autonomous, traversable boundary-layer probe for flight testing on large aircraft is presented. The device is an innovative approach to conduct time-resolved measurements of boundary-layers with high resolution and accuracy. The probe can be installed quickly and at a wide range of positions on an aircraft. It consists of two modules. The first houses a movable arm with three fast response differential pressure transducers in the tip combined to a three-hole probe to measure dynamic pressure and flow direction, or longitudinal and transversal velocities, respectively, comprising both, mean values and fluctuations. The second module houses the main electronics and the power supply. The two modules are connected by a cable; communication between the probes and a control computer inside the aircraft cabin is established via radio. Neither cable connection between the device and the aircraft is needed, not any intrusion of the aircraft’s surface. Four probes of this type were in use during in-flight measurements with the “Advanced Technology Research Aircraft” Airbus A320-232 “D-ATRA” of the German national aeronautics and space research center “DLR”. Herein, the general design and the means to calibrate and to validate the reliability of the device are presented, along with some experiences and results from flight testing as examples.

This study has investigated the use of Dynamic Pressure-Sensitive Paint (DPSP) to analyse the complex multi-dimensional flow-field within a generic small-scale rectangular planform cavity. The analysis and understanding of cavity flows have received renewed interest due to the low-observability requirements for modern military aircraft and Unmanned Combat Aerial Vehicles (UCAVs). This study has demonstrated the use of DPSP to provide unsteady pressure data, from the full model surface coverage of a model-scale cavity ceiling in a transonic Mach number range of 0.70 ≤ M ≤ 0.95. These results were compared against traditional point transducer measurements and also compared with numerical simulations of the cavity flow-field. These comparisons demonstrate that DPSP represents a very useful tool for the validation of numerical data as well as for flow-diagnostic investigations, where point measurements from transducers would provide spatially limited data. Further analysis of the transducer and DPSP data has also demonstrated temporal intermittency of the widely known Rossiter modes within the cavity acoustic signature. This instability has been previously linked to mode switching and a breakdown in the modal generation feedback process within the cavity flowfield.

A rapid assessment approach based on oblique-shock-wave and isentropic relations is established to analyze multiple freestream parameter matching schemes in ground-based tests of airframe/propulsion integrated vehicles in a combustion wind tunnel. Shock and isentropic compression configurations are introduced to estimate the deviation range of aerodynamic properties to ture flight. It is found that the deviation of the oblique shock angle and the post-shock pressure coefficient between in vitiated and clean air is rather small while for an isentropic compression configuration the pressure coefficient deviation is comparatively big, when the incoming flow Mach number is matched. If the total pressure and dynamic pressure are simultaneously matched regardless of the restrictions on the matching of Mach number, however, a remarkable reduce emerges in the difference of wave structures as well as aerodynamic forces. Analysis shows that it is the decrease of incoming flow Mach number that make the P0Q-matched condition better than the M-matched condition. Considering the aero-propulsion integrated performance, h0P0Q matching scheme is an instructive choice. Numerical simulation further indicates that if appropriate flow parameter matching scheme is selected, combustion wind tunnels can even reproduce the aerodynamic characteristics closer to true flight than conventional ones.

This describes the evolution of a high pressure combustion rig with viewing capability from 3 dimensions to add functionality for hypersonic testing. The design target is for supersonic combustion and hypersonic flow characterisation. One of the current major features of the existing combustion rig is the ability to carry out full 3D studies via the use of an extra optical window which allows flow and other parameter characterization in the rθ direction. Non intrusive laser diagnostics have been installed to use this, with the analysis shown with dynamic pressure transducers to give thermoacoustic relevance to the heat release detection capability.

In hypersonic flight, sharp body concepts have several advantages compared to the classic blunt approach. Conversely, the heat load on the sharp leading edge is significantly increased. Therefore, it is necessary to consider active cooling concepts for the sharp leading edge. In this paper a transpiration-cooled cone is investigated with focus on the varying cooling mass flow along the surface, caused by anisotropic permeability of the material, and increasing wall thickness. Two cones with different angles are investigated with respect to their cooling flow behavior in the DLR test facility AORTA, which was built to characterize porous material. The measurements include the dynamic pressure distribution at the surface, in addition to reservoir pressure and the total mass flow. Furthermore a numerical approach is presented to simulate the pressure loss over the wall thickness with the two-dimensional Darcy equation and the mass flow distribution at the surface. Finally, the results from the experiment and the simulation are presented and assessed with respect to the fiber orientation of the material. The results show clear trends of mass flow distribution along the circumference of the cone. These mass flow distributions can be explained by the flow dependence of the fiber orientation and the resulting permeability.

A Model Test Apparatus (MTA) to support and control wind tunnel models was developed and installed at the Air Force Institute of Technology’s low speed wind tunnel. The MTA is a 6-actuator robotic manipulator arm which was shown to be capable of dynamic motions. Maximum accelerations of 1G were performed in all three axes, as well as linear velocities in excess of 3 m/s. Maximum rotational rates of 2π rad/s were demonstrated. The MTA was installed so as to position and manipulate models through an access port on the side wall of the Air Force Institute of Technology low speed tunnel. This paper describes the initial development work to integrate the MTA with the tunnel, dictate model motion, and gather data from dynamically moving models. Initial aerodynamic studies were carried out for low reduced-frequency (k = 0.001 to 0.002) oscillatory-pitch motions for a rectangular wing with an aspect ratio of 2 and a NACA 0012 airfoil section. Dynamic pressure data indicated the abrupt stall, consistent with the literature on the subject.

The major disturbance to the launch vehicle during the high dynamic pressure period is inflight winds, and it causes extra aerodynamic load and significant magnitude of bending moment. Load relief system is used to decrease the aerodynamic load by reducing the angle of attack to increase the launching capability. This paper focuses on the load control technology when the launch vehicle passes through the strong wind area. Firstly, by analyzing and simulating the load relief system of the traditional launch vehicles including long march 3A, atlas/centaur, Saturn V, titan III and Ariane 5, this paper points out the advantages and disadvantages of classical passive trajectory correction method and four active load control methods. Besides, the control enhanced and parameterized technology (CEPT) is putted forward from the active disturbance rejection control (ADRC) and load relief cell. On the one hand, the error signal observed by the modified state observer is introduced to compensate the uncertain error signal observed by extended state observer in the ADRC loop; on the other hand, the load relief cell is introduced into the initial control system to enhance the load relief capability; at last, the control law of the CEPT controller is parameterized, a method is given to select the gains in the control law, and the load restriction effect under the combination of CEPT and initial controller is shown. Taking a kind of liquid launch vehicle as an example, and considering the inherent elastic vibration and liquid sloshing, this paper compares the load relief effect under different control architectures motioned above. The simulation result indicates that the CEPT controller expands the range of the parameters, enhances the disturbance rejection capacity for the flight control system, and efficiently improves the load control effect.

Jet noise has been an active area of research owing to public health concerns and aviation regulations. This problem is further amplified with the use of multiple engines placed in closed proximity to each other. In addition to significantly altered farfield noise as compared to a single-jet, previous experimental results show high levels of dynamic pressure fluctuations in the inter-nozzle region which can cause structural damage. In this work, Large Eddy Simulations are performed to study the dynamics of a supersonic twinjet configuration, with and without flow control. Results indicate that relative to single-jet configuration, twinjets show decrease in column lengths due to significant modifications in shear-layer properties. The simulations also corroborate previous experimental observations of very high levels of dynamic pressure fluctuations in the inter-nozzle region relative to a single-jet configuration for identical flow parameters. The phenomenon of noise shielding along the plane containing the two jets has been reported for all the twinjet cases explored in this work and trends of sound pressure levels obtained at various polar angles indicate an increased efficiency of shielding at lower polar angles. The sensitivity of the dynamic pressure fluctuations on the nozzle-exit boundary layer thickness is also characterized. Simulations are also performed to mimic the effects of Localized Arc Filament Plasma Actuators at different excitation frequencies (St=0.3 and 0.9) actuated in an axisymmetric fashion (m=0). Close to the nozzle exit, toroidal structures formed as a result of control result in significant increase in jet spreading leading to the formation of a recirculation region along the symmetry plane, relative to the uncontrolled case. This excitation scheme increased the levels of dynamic pressure fluctuations closer to the nozzle exit in the inter-nozzle region. In contrast, the near field shows significant reduction in pressure fluctuations, especially for the higher Strouhal number case.

Accurate prediction of wall-bounded flows with Reynolds-averaged Navier-Stokes turbulence models requires the improved modeling of physical processes that occur in turbulent flows such as the interaction of turbulent velocity and pressure fields, turbulent diffusion, and dissipative processes. Recently, novel models for the velocity/pressuregradient correlations through the fourth order were proposed by the authors and successfully validated against direct numerical simulation (DNS) data in wall-bounded flows without separation in a priori testing. In the current work, these models are implemented in transport equations for the Reynolds stresses to evaluate the model performance in a fullydeveloped planar channel. Other terms that need modeling in the Reynolds stress transport equations are represented in simulations by their DNS profiles to eliminate uncertainty due to modeling these terms. Problems with using DNS data as an input in simulations are highlighted. An approach to mitigate some of those problems is proposed. The computed results are compared against DNS data.

The NASA Low Density Supersonic Decelerators (LDSD) project includes a highaltitude Supersonic Flight Dynamics Test (SFDT) that tests aerodynamic decelerators in the correct Mach/dynamic pressure environment for use on Mars landing missions. The 30.5 meter parachute of LDSD was developed to be mortar-deployed for future Mars use. However, the unique design of the Supersonic Flight Dynamics Test, with a center-mounted solid rocket, and the mass of the parachute (100 kg) precluded the use of a CG-aligned parachute mortar and required the development of a staged pilot deployment that preserved the canopy extraction characteristics of a mortar-deployed parachute. The architecture of the parachute deployment is discussed, including the use of a stiffened triple bridle, the elimination of the metallic bridle confluence fitting, the use of two deployment bags and the separation of the pilot device prior to line stretch. The extensibility to a mortar-deployed parachute is discussed along with the parameters that were assumed for a successful deployment and inflation. A multi-body dynamic simulation was developed to predict the performance of the parachute deployment and to tune the deployment parameters. The results of the simulation are compared to the performance of the system in the first SFDT flight, as well as ground-based development tests of the deployment sequence.

During the first Supersonic Flight Dynamics Test (SFDT-1) for NASA's Low Density Supersonic Decelerator (LDSD) Program, the Parachute Decelerator System (PDS) was successfully tested. The main parachute in the PDS was a 30.5-meter supersonic Disksail parachute. The term Disksail is derived from the canopy's constructional geometry, as it combined the aspects of a ringsail and a flat circular round (disk) canopy. The crown area of the canopy contained the disk feature, as a large flat circular disk that extended from the canopy's vent down to the upper gap. From this upper gap to the skirt-band the canopy was constructed with characteristics of sails seen in a ringsail. There was a second lower gap present in this sail region. The canopy maintained a nearly 10x forebody diameter trailing distance with 1.7 Do suspension line lengths. During the test, the parachute was deployed at the targeted Mach and dynamic pressure. Although the supersonic Disksail parachute experienced an anomaly during the inflation process, the system was tested successfully in the environment it was designed to operate within. The nature of the failure seen originated in the disk portion of the canopy. High-speed and high-resolution imagery of the anomaly was captured and has been used to aid in the forensics of the failure cause. In addition to the imagery, an inertial measurement unit (IMU) recorded test vehicle dynamics and loadcells captured the bridle termination forces. In reviewing the imagery and load data a number of hypothesizes have been generated in an attempt to explain the cause of the anomaly.

An inflatable supersonic decelerator that enables p ayloads larger than the Mars Science Lab to be delivered to the surface of Mars has been developed. This decelerator, the SIAD- R, underwent a full-scale suborbital supersonic tes t flight in 2014. During that flight, the SIAD-R deployed at roughly Mach 4.1 at a dynamic pressure of 350 Pa. The SIAD-R performed flawlessly throughout this test flight an d has achieved Technology Readiness Level 6. This paper describes the test flight envir onment and the response of the SIAD-R to the test flight environment. Specifically, characte rization of the SIAD-R deployment, the inflation pressure of the SIAD-R, and the aeroelast ic distortion of the SIAD-R were measured and recorded during the flight and are discussed. Notably, the approximate proportionality of SIAD-R aeroelastic distortion to the ratio of dynamic pressure to SIAD-R inflation pressure that had been previously identif ied in subsonic testing was also found to be valid during in this supersonic test flight; this a pproximation allows for extrapolation of the aeroelastic distortion under a dynamic pressure of 2200 Pa even though this dynamic pressure was not achieved during the 2014 test flig ht. This paper also discusses the current design of the SIAD-R, with the emphasis on the inte rfaces to the supersonic test flight vehicle and design updates that have been made to the SIAD-R following the completion of subsonic testing and vacuum chamber inflation testing.

The paper presents a comprehensive study involving experiments and numerical simulations to measure the acoustic field in a rectangular supersonic cavity with and without a generic store. Unsteady pressure spectra measured in the empty cavity show the presence of highamplitude, discrete tones at all measurement locations with highest fluctuation levels at the downsteam wall of the cavity. The frequency of cavity tones is well predicted by Rossiter modes of resonance. Pressure fluctuation levels increase with Reynolds number and assumes a universal character when non-dimensionalized with freestream dynamic pressure. In the presence of a store, not only the resonance modes in the cavity are altered but the energy is redistributed among various modes. A reasonable correlation is observed between CFD and experimental data.

This paper will describe the evolution of a high pressure optical gas turbine rig with viewing capability from 3 dimensions. Commisioning work is described followed by an introduction to a generic swirl burner, scaled up by 40 % from a well characterised atmospheric pressure unit. One of the major features of the system is the ability to carry out full 3D studies via the use of an extra optical window which allows flow and other parameter characterization in the rθ direction. Non intrusive laser diagnostics have been installed to use this, with the analysis shown with dynamic pressure transducers to give thermoacoustic relevance to the heat release detection capability.

An uncertainty-based simulation work flow is used to automate the prediction of steadystate aerodynamic loads for a design variant of the HIFiRE-6 hypersonic flight research vehicle. Tools distributed with the Chimera Grid Tools (CGT) framework are exercised to perform overset mesh generation and verify domain connectivity. OVERFLOW is used to compute Reynolds-Averaged Navier-Stokes estimates of static aerodynamic loads at different points within a design space representative of the vehicle’s planned flight test environment. Sensitivity analysis of a design space composed of Mach number, angle of attack, sideslip angle, and dynamic pressure is performed to assess the suitability of using secondorder central-composite statistical designs to capture modest variations in customer-defined system response quantities. Within the specified bounds, automated simulation analytics suggest that for the fixed vehicle being analyzed, variations in axial and normal force coefficients are driven by perturbations in angle of attack and Mach number. Perturbations in sideslip angle induce the most variation in side force coefficient. Perturbations in an interaction term involving sideslip angle and angle of attack, and sideslip angle are the primary drivers impacting variation in rolling moment coefficient. Perturbations in angle of attack and Mach number are predicted to be the primary drivers of variations in pitching moment coefficient. Perturbations in sideslip angle and Mach number are predicted to be the primary drivers for variations in yawing moment. Perturbations in dynamic pressure were determined to be statistically insignificant for the design space analyzed. Analysis of variation results of the design indicate that second-order surrogates, based on a small number of independent simulations, are able to capture 90 to 100% of the variations in steady-state aerodynamic loads predicted by simulation associated with a vehicle at a fixed point along its trajectory. We expect these automated techniques and observations to be useful for future design, testing, and evaluation activities involving high-speed vehicles.

The dynamics of blunt bodies under parachutes is complex. The designer must balance translational requirements to slow down rapidly and over short distance, without violating rotational requirements to avoid high rates, flipover, and riser recontact. The designer’s effectors are: parachute drag area, the timing of reefing stages, timing of chute cutaway, and timing of next stage deployment. In the time domain, the equations of motion are non-linear and time varying, which has limited their solution to the most idealized cases, among them the steady state flight at constant dynamic pressure. The steady state dynamic pressure assumption is quite limiting as the whole purpose of parachutes is to reduce the dynamic pressure in a series of highly transient events. This paper explores the body/chute system dynamics with distance as an independent variable, rather than time. We find in the spatial coordinates, the differential equation for dynamic pressure becomes linear and yields a simple closed form solution. But the main finding here is that translational dynamics alter the rotational stability during the velocity transient period. This effect is a strong function of the difference in drag area between the previous stage and the next, and decays to zero as the system approaches steady state velocity. The transient term is found to be stabilizing at cutaway (reducing drag area) and destabilizing for reefed systems (increasing drag area).

Conical shell theory and a supersonic potential flow aerodynamic theory are used to study the nonlinear pressure buckling and aeroelastic limit cycle behavior of the thermal protection system for NASA's Hypersonic Inflatable Aerodynamic Decelerator. The structural model of the thermal protection system consists of an orthotropic conical shell of the Donnell type, resting on several circumferential elastic supports. Classical Piston Theory is used initially for the aerodynamic pressure, but was found to be insufficient at low supersonic Mach numbers. Transform methods are applied to the convected wave equation for potential flow, and a time-dependent aerodynamic pressure correction factor is obtained. The Lagrangian of the shell system is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the governing differential-algebraic equations of motion. Aeroelastic limit cycle oscillations and buckling deformations are calculated in the time domain using a Runge-Kutta method in MATLAB. Three conical shell geometries were considered in the present analysis: a 3-meter diameter 70 deg. cone, a 3.7-meter 70 deg. cone, and a 6-meter diameter 70 deg. cone. The 6-meter configuration was loaded statically and the results were compared with an experimental load test of a 6-meter HIAD. Though agreement between theoretical and experimental strains was poor, the circumferential wrinkling phenomena observed during the experiments was captured by the theory and axial deformations were qualitatively similar in shape. With Piston Theory aerodynamics, the nonlinear flutter dynamic pressures of the 3-meter configuration were in agreement with the values calculated using linear theory, and the limit cycle amplitudes were generally on the order of the shell thickness. The effect of axial tension was studied for this configuration, and increasing tension was found to decrease the limit cycle amplitudes when the circumferential elastic supports were neglected, but resulted in more complex behavior when the supports were included. The nominal flutter dynamic pressure of the 3.7-meter configuration was significantly lower than that of the 3-meter, and it was found that two sets of natural modes coalesce to flutter modes near the same dynamic pressure. This resulted in a significant drop in the limit cycle frequencies at higher dynamic pressures, where the flutter mode with the lower frequency becomes more critical. Pre-buckling pressure loads and the aerodynamic pressure correction factor were studied for all geometries, and these effects resulted in significantly lower flutter boundaries compared with Piston Theory alone. The maximum dynamic pressure predicted by aerodynamic simulations of a proposed 3.7-meter HIAD vehicle was still lower than any of the calculated flutter dynamic pressures, suggesting that aeroelastic effects for this vehicle are of little concern.

A study of the flutter characteristics of three different wing planforms representative of present and future generations of fighters is presented. The main objective is to obtain a better understanding of complex transonic flutter behaviors involving LCOs triggered by natural mode instabilities at low dynamic pressures, followed by transitions to strong bending-torsion or multi-mode flutter at higher dynamic pressures. Two of the wings have flutter boundaries with deep "chimneys" in the vicinity of Mach 1. Inside the chimneys, the LCO amplitudes are insensitive to dynamic pressure or density and persist down to very low densities representative of altitudes in the high stratosphere. The Mach number freeze or transonic stabilization phenomenon plays an important role in determining the shape of the SDOF and chimney regions. Near the freeze Mach number, the unsteady aerodynamic forces show rapid reversals in amplitude and phase, which have a profound influence on the flutter boundary and the observed flutter characteristics. The transonic stabilization mechanism appears responsible for the rapid quenching of natural mode LCOs and the creation of deep chimneys in the flutter boundary near Mach 1, as observed in several wind tunnel tests.

In this work we propose a new approach to elucidate specific combustion instability driving mechanisms and demonstrate application of the developed method to data collected using modern diagnostics. A laboratory-scale, optically accessible combustor was developed that exhibits full-scale-like, transverse combustion instabilities. We acquired simultaneous high-repetition rate (10 kHz) diagnostic data using particle image velocimetry, OH-planar laser induced fluorescence, CH-chemiluminescence, and dynamic pressure measurements. This paper discuses our approach to analyze the many gigabytes of data collected per second of testing. Specifically, we developed a technique based on the dynamic mode decomposition (DMD) and used it to extract interesting spatial and temporal features of the near-injector region from the simultaneously acquired data. This developed technique was used to analyze two datasets: one taken with the rig operating in a stable configuration and the other obtained when the rig was experiencing a transverse combustion instability. The analysis showed that both sets of data exhibited similar hydrodynamic and combustion modes/oscillations, suggesting that the combustor stability depends primarily on the phase of the of the energy transfer from the unsteady combustion heat release to the natural acoustic modes of the combustor.