SEARCH · Engineering Papers
Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The effect of Orbiter flight control on tether dynamics is studied by simulation. Open-loop effects of Orbiter jet firing on tether dynamics are shown, and the potential for closed-loop interaction between tether dynamics and Orbiter flight control is determined. The significance of these effects on Orbiter flight control and tether control is assessed.
Pretest predictions of the dynamic characteristics of the XV-15 tilting proprotor aircraft are presented. The data for the aircraft in flight include: trim conditions, flight dynamics, gust response, aeroelastic stability and the wing response to control. The data for the aircraft in the Ames 40- by 80-ft wind tunnel include aeroelastic stability and the wing response to control. The calculations were made for pylon tilt angles of 0 deg (airplane mode), 0.1 deg (pylon unlocked), 30, 60, and 90 deg (helicopter mode).
Experiments as well as accompanying simulations are described that serve in preparation of a space flight experiment to study the dynamics of miscible interfaces. The investigation specifically addresses the importance of both nonsolenoidal effects as well as nonconventional Korteweg stresses in flows that give rise to steep but finite concentration gradients. The investigation focuses on the flow in which a less viscous fluid displaces one of higher viscosity and different density within a narrow capillary tube. The fluids are miscible in all proportions. An intruding finger forms that occupies a fraction of the total tube diameter. Depending on the flow conditions, as expressed by the Peclet number, a dimensionless viscosity ratio, and a gravity parameter, this fraction can vary between approximately 0.9 and 0.2. For large Pe values, a quasi-steady finger forms, which persists for a time of O(Pe) before it starts to decay, and Poiseuille flow and Taylor dispersion are approached asymptotically. Depending on the specific flow conditions, we observe a variety of topologically different streamline patterns, among them some that leak fluid from the finger tip. For small Pe values, the flow decays from the start and asymptotically reaches Taylor dispersion after a time of O(Pe). Comparisons between experiments and numerical simulations based on the 'conventional' assumption of solenoidal velocity fields and without Korteweg stresses yield poor agreement as far as the Pe value is concerned that distinguishes these two regimes. As one possibility, we attribute this lack of agreement to the disregard of these terms. An attempt is made to use scaling arguments in order to evaluate the importance of the Korteweg stresses and of the assumption of solenoidality. While these effects should be strongest in absolute terms when steep concentration fronts exist, i.e., at large Pe, they may be relatively most important at lower values of Pe. We subsequently compare these conventional simulations to more complete simulations that account for nonvanishing divergence as well as Korteweg stresses. While the exact value of the relevant stress coefficients are not known, ballpark numbers do exist, and their use in the simulations indicates that these stresses may indeed be important. We plan to evaluate these issues in detail by means of comparing a space experiment with corresponding simulations, in order to extract more accurate Korteweg stress coefficients, and to confirm or deny the importance of such stresses.
We propose a space flight experiment to study the dynamics of miscible interfaces. A less viscous fluid displaces one of higher viscosity within a tube. The two fluids are miscible in all proportions. An intruding "finger" forms that occupies a fraction of the tube. As time progresses diffusion at the interface combined with flow induced straining between the two fluids modifies the concentration and velocity distributions within the whole tube. Also, under such circumstances it has been proposed that the interfacial stresses could depend on the local concentration gradients (Korteweg stresses) and that the divergence of the velocity need not be zero, even though the flow is incompressible. We have obtained reasonable agreement for the tip velocity between numerical simulations (that ignored the Korteweg stress and divergence effects) and physical experiments only at high Peclet Numbers. However at moderate to low Pe agreement was poor. As one possibility we attributed this lack of agreement to the disregard of these effects. We propose a space experiment to measure the finger shape, tip velocity, and the velocity and concentration fields. From intercomparisons between the experiment and the calculations we can then extract values for the coefficients of the Korteweg stress terms and confirm or deny the importance of these stresses.
We propose a space flight experiment to study the dynamics of miscible interfaces. A less viscous fluid displaces one of higher viscosity within a tube. The two fluids are miscible in all proportions. An intruding "finger" forms that occupies a fraction of the tube. As time progresses diffusion at the interface combined with flow induced straining between the two fluids modifies the concentration and velocity distributions within the whole tube. Also, under such circumstances it has been proposed that the interfacial stresses could depend on the local concentration gradients (Korteweg stresses) and that the divergence of the velocity need not be zero, even though the flow is incompressible. We have obtained reasonable agreement for the tip velocity between numerical simulations (that ignored the Korteweg stress and divergence effects) and physical experiments only at high Peelet Numbers. However at moderate to low Pe agreement was poor. As one possibility we attributed this lack of agreement to the disregard of these effects. We propose a space experiment to measure the finger shape, tip velocity, and the velocity and concentration fields. From intercomparisons between the experiment and the calculations we can then extract values for the coefficients of the Korteweg stress terms and confirm or deny the importance of these stresses.
The Wideband Analog System (WADS) which performs post flight structural dynamic analysis of wideband FM data recorded during orbital flight tests is defined. Screening, reporting, and plotting steady state and transient wave analysis data are provided by WADS.
Supersonic inflatable aerodynamic decelerators (IADs) are drag devices intended to be deployed at high Mach numbers. In the application considered here they assist in the descent and landing of spacecraft on Mars. Although promising, present IAD technology is not yet sufficiently mature for use in the near future. This paper describes a technology maturation plan for tension cone IADs using subscale test articles to reduce development costs. As envisioned, the proposed test plan includes three phases: wind tunnel tests (subsonic), unpowered high-altitude flight tests (transonic), and powered high-altitude tests (supersonic). This test plan is based on a building block approach in which successful completion of each phase adds to the understanding of the behavior of IADs and reduces the risk of the subsequent, more expensive phases. By properly scaling the IADs, test articles of the same size and nearly the same construction can be used for all three phases. The final phase is a dynamically scaled flight test with IAD deployment at the same Mach number as the full-scale vehicle on Mars. Two full-scale example cases are presented: one for a single-stage system (15 m dia. IAD to subsonic retropropulsion), and another for a two-stage system (10.5 m dia. IAD to subsonic parachute). Using scale factors of 0.333 and 0.476 yield subscale test IADs of 5 m dia. The dynamically scaled powered flight test starts at Mach 4 and an altitude of 33.5 km. Existing balloons and rocket motors are shown to be adequate to meet the required test conditions.
The Dragonfly mission will place a rotorcraft/lander on Titan by 2034. The entry, descent, and landing system of the Dragonfly mission includes two parachutes: a drogue and a main. To provide needed data, subscale drop tests were used to conduct an experimental investigation of the aeroshell/parachute dynamics. The drop tests used a Disk-Gap-Band drogue parachute and two types of Ringslot main parachutes. All tests used a representative aeroshell which included the heatshield. All models were geometrically scaled to 16.7 percent. The model aeroshell had a diameter of 0.75 m. The model parachute nominal diameters were 0.9 m for the drogue and 2.78 m for the main. The aeroshell’s mass properties were dynamically scaled to simulate flight at an altitude of 4 km at Titan. This dynamic scaling allowed the conversion of model test results to full-scale Titan conditions. Onboard instrumentation on the aeroshell provided data on the rotation rates, from which the Euler angles were determined. Tests were conducted by lifting the models with a drone to an altitude of 350 m and dropping them inverted. Key results from these tests were: 1) the models were able to recover from the extreme inverted initial condition and settle to low-amplitude oscillations; 2) ninety nine percent of the time the oscillation amplitudes observed with the drogue parachute were 11.4 degrees or less; 3) ninety nine percent of the time the oscillation amplitudes observed with the 20 percent porosity main parachute were 15.4 degrees or less.
The Dragonfly mission will place a rotorcraft/lander on Titan by 2034. The entry, descent, and landing system of the Dragonfly mission includes two parachutes: a drogue and a main. To provide needed data, subscale drop tests were used to conduct an experimental investigation of the aeroshell/parachute dynamics. The drop tests used a Disk-Gap-Band drogue parachute and two types of Ringslot main parachutes. All tests used a representative aeroshell which included the heatshield. All models were geometrically scaled to 16.7 percent. The model aeroshell had a diameter of 0.75 m. The model parachute nominal diameters were 0.9 m for the drogue and 2.78 m for the main. The aeroshell’s mass properties were dynamically scaled to simulate flight at an altitude of 4 km at Titan. This dynamic scaling allowed the conversion of model test results to full-scale Titan conditions. Onboard instrumentation on the aeroshell provided data on the rotation rates, from which the Euler angles were determined. Tests were conducted by lifting the models with a drone to an altitude of 350 m and dropping them inverted. Key results from these tests were: 1) the models were able to recover from the extreme inverted initial condition and settle to low-amplitude oscillations; 2) ninety nine percent of the time the oscillation amplitudes observed with the drogue parachute were 11.4 degrees or less; 3) ninety nine percent of the time the oscillation amplitudes observed with the 20 percent porosity main parachute were 15.4 degrees or less.
Pilot performance was studied during simulated manual terrain following flight for ride quality criteria validation. An existing B-1 simulation program provided the data for these investigations. The B-1 simulation program included terrain following flights under varying controlled conditions of turbulence, terrain, mission length, and system dynamics. The flight simulator consisted of a moving base cockpit which reproduced motions due to turbulence and control inputs. The B-1 aircraft dynamics were programmed with six-degrees-of-freedom equations of motion with three symmetric and two antisymmetric structural degrees of freedom. The results provided preliminary validation of existing ride quality criteria and identified several ride quality/handling quality parameters which may be of value in future ride quality/criteria development.
The basic ability of STARS to maintain a satellite communications link with TDRSS satellites during dynamic aircraft flights was successfully demonstrated during FD 1. The Range Safety and Range User systems' link margins were measured. The ability to acquire/reacquire and maintain lock between a high-dynamic vehicle and a satellite-based system was demonstrated. The Range Safety system simultaneously received and processed command links from space and ground transmitters and provided near real-time Range Safety telemetry to DFRC, which then sent it in near real time to KSC, GSFC, and WFF for monitoring. The GPS receiver maintained track except during extremely dynamic maneuvers. The Range User system sent data at three different data rates. There were excellent cooperation and support from the different Centers, contractors, and Ranges. A large amount of data was recorded and extensive post-flight analysis was performed. The Range User TDRSS link margin met or exceeded the predicted performance at three different data rates. The Range Safety launch-head link margins generally agreed with the predicted performance. The UPS positions and velocities agreed with those from tracking radar to within about 20 m and a few rn/s. The link margins for the Range Safety TDRSS telemetry link were less than expected. The link margin for one TDRSS command link LPT channel was occasionally much less than the other. Additional post-flight testing has yet to identify the root causes of these results. There were many lessons learned from this first set of test flights. The most important one is that more time and testing are needed for each step to deal with the inevitable problems. It is vital that these lessons be among the primary areas of study that will carry over from FD#1 to FD#2, which is currently scheduled for early FY05 at DFRC and will use a specially designed Ku-band phased array antenna for the Range User system. The next series of flight demonstrations scheduled for late 2004 at DFRC will incorporate many lessons learned from FD#1. A specially designed Ku-band phased array antenna will be used with the Range User system. A test flight on a hypersonic vehicle is planned by the end of 2006.
The X-29A advanced technology demonstrator flight envelope expansion program and the subsequent flight research phase gave impetus to the development of several innovative real-time analysis and display techniques. These new techniques produced significant improvements in flight test productivity, flight research capabilities, and flight safety. These techniques include real-time measurement and display of in-flight structural loads, dynamic structural mode frequency and damping, flight control system dynamic stability and control response, aeroperformance drag polars, and aircraft specific excess power. Several of these analysis techniques also provided for direct comparisons of flight-measured results with analytical predictions. The aeroperformance technique was made possible by the concurrent development of a new simplified in-flight net thrust computation method. To achieve these levels of on-line flight test analysis, integration of ground and airborne systems was required. The capability of NASA Ames Research Center, Dryden Flight Research Facility's Western Aeronautical Test Range was a key factor in enabling implementation of these methods.
The X-29A advanced technology demonstrator flight envelope expansion program and the subsequent flight research phase gave impetus to the development of several innovative real-time analysis and display techniques. These new techniques produced significant improvements in flight test productivity, flight research capabilities, and flight safety. These techniques include real-time measurement and display of in-flight structural loads, dynamic structural mode frequency and damping, flight control system dynamic stability and control response, aeroperformance drag polars, and aircraft specific excess power. Several of these analysis techniques also provided for direct comparisons of flight-measured results with analytical predictions. The aeroperformance technique was made possible by the concurrent development of a new simplified in-flight net thrust computation method. To achieve these levels of on-line flight test analysis, integration of ground and airborne systems was required. The capability of NASA Ames Research Center, Dryden Flight Research Facility's Western Aeronautical Test Range was a key factor to enable implementation of these methods.
Advantages and disadvantages of using simulation languages in solving structural problems are given. Structural and solid mechanics problems which have strong interactions with other disciplines are emphasized. An aeroservoelastic illustration is described with significant interactions between the dynamics of a flexible flight vehicle structure, the aerodynamics to which it is subjected, the dynamic flight equations, and the vehicle's servo-control system.
The role of the system error dynamics in model-following control systems is discussed, along with the use of handling quality specifications, actuation bandwidth constraints, stability, and closed-loop performance requirements in flight control design. The model-following problem is formulated using both direct state-space and linear, quadratic, optimization techniques. The results are then demonstrated using several examples involving a generic forward-swept-wing vehicle and a conventional flight vehicle with large parameter uncertainty in order to illustrate the trade-off in closed-loop performance and control law complexity.
A flight test program was undertaken to demonstrate the feasibility of obtaining external burning (EB) data in a flight environment, and to address the question of facility interference in ground test external burning data. Results showed that external burning was effective at reducing transonic base drag in a flight environment having dynamically changing conditions. The flight test program demonstrated that external burning is not just a laboratory phenomenon but is a viable technology for aerospace vehicles.
The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.
Shock spectra are presented for the TC-1 (Viking dynamic simulator) flight.