Earth Entry Research Articles

Products from NASA's in-space propulsion technology program applicable to low-cost planetary missions

Since September 2001, NASA's In-Space Propulsion Technology (ISPT) program has been developing technologies for lowering the cost of planetary science missions. Recently completed is the high-temperature Advanced Material Bipropellant Rocket (AMBR) engine providing higher performance for lower cost. Two other cost saving technologies nearing completion are the NEXT ion thruster and the Aerocapture technology project. Under development are several technologies for low-cost sample return missions. These include a low-cost Hall-effect thruster (HIVHAC) which will be completed in 2011, light-weight propellant tanks, and a Multi-Mission Earth Entry Vehicle (MMEEV). This paper will discuss the status of the technology development, the cost savings or performance benefits, and applicability of these in-space propulsion technologies to NASA's future Discovery, and New Frontiers missions, as well as their relevance for sample return missions.

Mars sample return mission architectures utilizing low thrust propulsion

The Mars sample return mission is a flagship mission within ESA's Aurora program and envisioned to take place in the timeframe of 2020–2025. Previous studies developed a mission architecture consisting of two elements, an orbiter and a lander, each utilizing chemical propulsion and a heavy launcher like Ariane 5 ECA. The lander transports an ascent vehicle to the surface of Mars. The orbiter performs a separate impulsive transfer to Mars, conducts a rendezvous in Mars orbit with the sample container, delivered by the ascent vehicle, and returns the samples back to Earth in a small Earth entry capsule. Because the launch of the heavy orbiter by Ariane 5 ECA makes an Earth swing by mandatory for the trans-Mars injection, its total mission time amounts to about 1460 days. The present study takes a fresh look at the subject and conducts a more general mission and system analysis of the space transportation elements including electric propulsion for the transfer. Therefore, detailed spacecraft models for orbiters, landers and ascent vehicles are developed. Based on that, trajectory calculations and optimizations of interplanetary transfers, Mars entries, descents and landings as well as Mars ascents are carried out. The results of the system analysis identified electric propulsion for the orbiter as most beneficial in terms of launch mass, leading to a reduction of launch vehicle requirements and enabling a launch by a Soyuz-Fregat into GTO. Such a sample return mission could be conducted within 1150–1250 days. Concerning the lander, a separate launch in combination with electric propulsion leads to a significant reduction of launch vehicle requirements, but also requires a large number of engines and correspondingly a large power system. Therefore, a lander performing a separate chemical transfer could possibly be more advantageous. Alternatively, a second possible mission architecture has been developed, requiring only one heavy launch vehicle (e.g., Proton). In that case the lander is transported piggyback by the electrically propelled orbiter.

Radiative-Collisional Models in Non-Equilibrium Aerothermodynamics of Entry Probes

To design space vehicles aimed for returning payloads from a geostationary orbit, the Moon and other large or small planets of Solar system, a knowledge of the total (convective and radiative) heating from an environment is required. It is well known that the radiative heat load on a space vehicle moving through the atmosphere increases as the speed and the size increase, therefore, in many of these missions the large part of the trajectory will pass at high altitude, where the low atmospheric density can lead to significant thermal, chemical and physical nonequilibrium effects. Physical models and computational codes used to predict the aerothermodynamics must account for not only high temperature equilibrium thermodynamics (as a rule, within the framework of the local thermodynamic equilibrium (LTE) approach), but also for nonequilibrium one. Therefore, an accurate prediction of radiative heating as well as convective one under both equilibrium and nonequilibrium conditions becomes important to designers and space mission planners. To develop a prediction computational fluid dynamics (CFD) tool for reentry flows, where dissociation, ionization and radiation are important, some major areas are addressed. The most significant of them are following: (1) physical-chemical kinetics of high temperature dissociated and ionized gases, (2) transport properties of the gas mixtures, (3) spectral radiation properties of high temperature gases and low-temperature plasmas, (4) numerical simulation algorithms for prediction of nonequilibrium gas mixtures dynamics and radiation heat transfer in volumes of various geometry, and (5) models of physical and chemical processes accompanied by interaction of gas flows and radiation with thermoprotection systems (TPS) of space vehicles (including their thermochemical destruction, ablation, sublimation, etc.). In literatures (See Refs. (Park, C, 1990, Nonequilibrium Hypersonic Aerothermodynamics, Willey-Interscience Publication, J. Wiley & Sons, New York; Park, C., 1993, “Review of Chemical Kinetic Problems of Future NASA Missions. I: Earth Entries,” J. Thermophys. Heat Transfer, 7(3), pp. 385–398; Park, et al., 1994, “Review of Chemical-Kinetic Problems of Future NASA Missions, II: Mars Entries,” J. Thermophys. Heat Transfer, 8(1), pp. 9–23; Sarma, G., 2000, “Physico-Chemical Modelling in Hypersonic Flow Simulation,” Prog. Aerosp. Sci., 36, pp. 281–349; Huo, and Thuemmel, 1995, Electron-Air Molecule Collisions in Hypersonic Flows. Molecular Physics and Hypersonic Flows, Capitelli M., ed., Kluwer Academic Publishers, pp. 115–138.)) one can find reviews of governing equations used in the aerophysics, boundary conditions and the associated inputs using the physical-chemical models and their partially successful applications. This article presents the states of the art of models of electronic kinetics in the nonequilibrium low-temperature plasma of complex chemical compositions (air and carbon dioxide mixtures) widely met in various aerospace applications. Special attention is given to electronic kinetics of atoms and diatomic molecules within the framework of the radiative-collisional models.

The effects of earthward directed interplanetary coronal mass ejections on near-Earth S band signal links

[1] Human space exploration is expected to enter its next phase in the coming decades as the United States prepares to return to the Moon or perhaps venture even further with a crewed mission to a near-Earth asteroid. Both mission classes are viewed by NASA as precursors of eventual crewed missions to Mars. In anticipation of extensive robotic and human presence in the space environment beyond the protection of the Earth's magnetosphere, it is important to better quantify and bound effects of earthward directed solar storms not just on the human body but also on engineering signals. In this paper, we study the effects of solar storms on S band (∼2.3 GHz) radio links in the near-Earth environment, primarily for application to navigation. In particular, we are concerned with induced long-period signatures on Doppler tracking data that could be confused with the Earth's gravity signature, resulting in perturbed trajectory solutions of returning spacecraft during Earth entry targeting. We have quantified “worst-case” levels of such induced signatures on S band signal phase using model predictions based on measured in situ charged particle content from satellites and have compared these results with signatures seen in actual tracking data during periods of interplanetary coronal mass ejections (ICME) and related geomagnetic storms. We show that induced Doppler can mask Earth gravity field effects in navigation trajectory solutions at S band, a commonly used frequency for near-Earth communications and navigation. Finally, we suggest a few ways that such effects can be identified, alleviated or eliminated in near real-time.

Automated Generation of Symmetric Lunar Free-Return Trajectories

A procedure for free-return trajectory generation in a simplified Earth-moon system is presented. With two-body and circular restricted three-body models, the algorithm constructs an initial guess of the translunar injection state and time of flight. Once the initial trajectory is found, the Jacobian of the constraints is derived analytically using linear perturbation theory, and a square system of nonlinear equations is solved numerically to target Earth entry interface conditions leading to feasible free-return trajectories. No trial and error is required to generate the initial guess. Possible free returns include departures from both posigrade and retrograde Earth orbits coupled with circumlunar or cislunar flight, in and out of the Earth-moon plane.

Initial Trajectory Model for a Multi-Maneuver Moon-to-Earth Abort Sequence

To support the mission design and trajectory design problems associated with the moon-to-Earth trajectories for the crew exploration vehicle, a starting trajectory model that serves as the first iterate for a complete targeting and optimization procedure that takes a spacecraft from any closed lunar parking orbit to the Earth entry interface state for any date is developed. The motivation for this work is to examine the any-time abort capability required for human moon missions. The results presented here are limited to impulsive maneuvers. An analytical procedure is developed that constructs a multi-impulse escape trajectory from the moon propagated forward in time and a backward propagated trajectory from the Earth with a mismatch in position and velocity near the sphere of influence of the moon. The position and velocity discontinuities at the mismatch point are small enough to lie within the convergence envelope of a simple gradient based differential correction procedure that can, at a minimum, generate a feasible solution. This solution can then be analyzed further and serve as an initial estimate for an optimization procedure. The efficiency of the method is illustrated by solving any-time abort transfer problems typical for a human mission.

Influence of Ablation on Radiative Heating for Earth Entry

Using the coupled ablation and radiation capability recently included in the LAURA flowfield solver, this paper investigates the influence of ablation on the shock-layer radiative heating for Earth entry. The extension of the HARA radiation model, which provides the radiation predictions in LAURA, to treat a gas consisting of the elements C, H, O, and N is discussed. It is shown that the absorption coefficient of air is increased with the introduction of the C and H elements. A simplified shock layer model is studied to show the impact of temperature, as well as the abundance of C and H, on the net absorption or emission from an ablation contaminated boundary layer. It is found that the ablation species reduce the radiative flux in the vacuum ultraviolet, through increased absorption, for all temperatures. However, in the infrared region of the spectrum, the ablation species increase the radiative flux, through strong emission, for temperatures above 3,000 K. Thus, depending on the temperature and abundance of ablation species, the contaminated boundary layer may either provide a net increase or decrease in the radiative flux reaching the wall. To assess the validity of the coupled ablation and radiation LAURA analysis, a previously analyzed Mars-return case (15.24 km/s), which contains significant ablation and radiation coupling, is studied. Exceptional agreement with previous viscous shock-layer results is obtained. A 40% decrease in the radiative flux is predicted for ablation rates equal to 20% of the free-stream mass flux. The Apollo 4 peak-heating case (10.24 km/s) is also studied. For ablation rates up to 3.4% of the free-stream mass flux, the radiative heating is reduced by up to 19%, while the convective heating is reduced by up to 87%. Good agreement with the Apollo 4 radiometer data is obtained by considering absorption in the radiometer cavity. For both the Mars return and the Apollo 4 cases, coupled radiation alone is found to reduce the radiative heating by 30 60% and the convective heating by less than 5%.

Ballute Entry Systems for Lunar Return and Low-Earth-Orbit Return Missions

This study investigates the feasibility of using ballutes for Earth entry at both lunarand low-Earth-orbit return velocities. Multiple entry strategies were investigated using analysis methods suitable for conceptual design and assuming anApollo-derived entry vehicle. Primary ballute size drivers are the thermal limitations andareal densities of the ballutematerial. Lunar-return entries that jettison the ballute after achieving low-Earth-orbit conditions were shown to reduce heating rates to within reusable thermal protection system limits. Deceleration was mitigated to approximately 4 g along a lunar-return trajectory when amoderate amount of lift was applied subsequent to ballute jettison. Investigation of a lower mass lunar-return cargo variant of the Crew Exploration Vehicle was shown to reduce ballute system mass and size. Ballute system mass as a percentage of the total entry mass was shown to be relatively independent of the entry mass for the range of lunar-return missions evaluated. However, material requirements indicate that continued technology development may be required for lunar-return system viability. Missions aimed at augmenting existing systems to provide downmass capability from the International Space Station showed promising results in terms of entry loads, requiredmaterial areal densities, and ballute systemmass and size.

Modeling of associative ionization reactions in hypersonic rarefied flows

When vehicles reenter the Earth’s atmosphere from space, the hypersonic conditions are sufficiently energetic to generate ionizing reactions. The production of a thin plasma layer around a hypersonic vehicle can block radio waves sent to and from the vehicle, leading to communications blackout. For Earth entry from orbit, the maximum energy involved in molecular collisions requires only associative ionization of air-species to be considered. In the present study, the modeling of such reactions is considered in detail using the direct simulation Monte Carlo (DSMC) method. For typical Earth entry conditions, with a velocity near 8km∕s, it is shown that the average ionizing reaction probabilities are small. Special numerical techniques must therefore be used in the DSMC technique in order to numerically resolve these reactions. Additional simulation problems arise from the relatively small mass of the electrons in comparison to the other atoms and molecules in these flow fields. Artificially increasing the electron mass greatly increases computational efficiency, and the viability of this approach is investigated. Simulation results are presented for conditions corresponding to the RAM-C II hypersonic flight experiment that gathered measurements of electron number density. It is demonstrated that simulation results for electron number density in this energy regime are relatively insensitive to the mass of the electrons. Direct comparison of DSMC results with the RAM-C II measurements for electron number density shows excellent agreement. These satisfactory comparisons represent the first direct verification of the ability of the DSMC technique to successfully predict the weak plasma generated around a hypersonic vehicle.

Abort determination with non-adaptive neural networks for the Mars precision landers

The 2009 Mars Science Laboratory (MSL) will attempt the first precision landing on Mars using a modified version of the Apollo Earth entry guidance program. The guidance routine, Entry Terminal Point Controller (ETPC), commands the deployment of a supersonic parachute while converging the range to the landing target. For very dispersed cases, ETPC is unlikely to converge the range to the target and command parachute deployment within Mach number and dynamic pressure constraints. A full-lift-up abort can save 85% of these failed trajectories while abandoning the precision landing objective. In order to implement an abort, a failed trajectory needs to be recognized in real time. The application of artificial neural networks (ANNs) as an abort determination technique was evaluated. An ANN was designed, trained and tested using Monte Carlo simulations of MSL descent for a severe dust storm scenario. When incorporated into ETPC, the ANN correctly classifies 87% of descent trajectories as abort or non-abort, reducing the probability of losing MSL in a severe dust storm from 18% to 3.5%. This research shows that ANNs are capable of recognizing failed descent trajectories and can significantly increase the survivability of MSL for very dispersed cases.

Direct Simulation Monte Carlo Simulations of Ballute Aerothermodynamics Under Hypersonic Rarefied Conditions

This paper presents computational results obtained with the direct simulation Monte Carlo method for towed ballute applications. A ballute is an inflatable drag device that can be used to create a large amount of drag at high altitudes. Consequently, ballutes provide a potential technology for achieving aerocapture when the primary spacecraft velocity reduction ( V) is achieved at much higher altitudes than with the conventional rigid aeroshell. Because the V is achieved at relatively high altitudes, rarefaction can be significant and is the motivation for the current study with the direct simulationMonte Carlo method. Computed surface and flowfield results are presented for a toroidal ballute, isolated tethers when exposed to freestream flow conditions, and the flow interactions resulting from a toroidal ballute when towed by a 6 m diameter Mars Pathfinder shaped (without tethers) spacecraft. All results presented are for Earth entry at velocities of 14 to 7 km=s (primary focus is at 8:55 km=s, same as some previous Titan aerocapture studies) and altitudes of 200 to 100 km. Variations of drag and heating coefficients as a function of rarefaction are presented. Adescription of theflow structure is provided and also an explanation of how it is affected by shock interactions produced solely by the ballute and those resulting from the two body combination of towed ballute and spacecraft is also given.

OXIDATION AND CATALYCITY OF THERMAL PROTECTION MATERIALS AT HIGH TEMPERATURE

When a space vehicle enters in planetary atmospheres like Earth (air) and/or Mars (CO2), oxidation and catalycity coupled phenomena occur on the thermal protective materials of the nose cap and wing leading edges conducting to an important excess of heating and possible damage of the materials. At the PROMES-CNRS laboratory (formerly IMP-CNRS), in the research field "Interaction high solar fluxes-matter", oxidation under plasma conditions, and particularly the active-passive transition in the oxidation of silicon-based ceramics and the catalytic recombination of atomic oxygen under the same conditions are studied. The most important conditions for the ground simulation (high temperature, non-equilibrium plasma flow) have been realized in the MESOX set-up which associates a solar radiation concentrator and a microwave generator. In this paper, are presented some results about the transition between the active and passive oxidation regimes and the recombination of atomic oxygen and/or carbon monoxide at the surface of sintered silicon carbide at high temperatures and in a total pressure range of 102-104 Pa. The comparison is made between the two atmospheres: air (Earth entry) and CO2 (Mars entry) for the oxidation transition domain and for the atoms recombination.

Planetary protection for Mars sample return

A Mars sample return (MSR) mission must deal with two new categories of planetary protection requirements: (1) assuring a very low probability of inadvertent release of the sample in order to provide extra protection against the extremely unlikely possibility of biological hazards in the returned material and (2) keeping the samples free of round-trip Earth organisms to facilitate confirmation of safety after return to Earth. This paper describes the planetary-protection-related technical challenges awaiting an MSR mission as well as the range of capabilities being developed to meet these challenges. Containment assurance requires breaking the chain of contact with Mars: the exterior of the sample container must not be contaminated with Mars material. The sample container and its seals must be designed to survive Earth impacts corresponding to the candidate mission profiles, and the Earth entry vehicle must be designed to withstand the thermal and structural rigors of Earth atmosphere entry (all with an unprecedented degree of confidence). Sample contamination must be avoided by sterilizing the entire spacecraft, a challenge with modern avionics, or by sterilizing the sample collection and containment gear and then isolating it from other parts of the spacecraft.

Sample Return Missions in the Coming Decades: A Comparison of Entry, Descent, and Landing Scenarios

In the coming decades, several missions will attempt to return samples to Earth from varying parts of the solar system. These samples will provide invaluable insight into the conditions present during the early formation of the solar system, and possibly give clues to how life began on Earth. A description of five sample return missions is presented (Stardust, Genesis, Muses-C, Mars Sample Return, and Comet Nucleus Sample Return). An overview of each sample return mission is given, concentrating particularly on the technical challenges posed during the Earth entry, descent, and landing phase of the missions. Each mission faces unique challenges in the design of an Earth entry capsule. The design of the entry capsule must address the aerodynamic, heating, deceleration, landing, and recovery requirements for the safe return of samples to Earth.

Review Mechanical properties of meteorites and their constituents

A review is presented of the mechanical properties of meteorites and meteorite constituents. Scientific literature data on the strength of stony and iron meteorites are extremely limited. The average mechanically-measured stony meteorite compressive strength is 200 MPa, while the average iron meteorite compressive strength is 430 MPa. However, the best current estimate of the strength of stony bodies in space may be in the range of only 1–5 MPa, based on observations of meteorite fragmentation due to dynamic atmospheric loading upon Earth entry. Mechanical property and behavior information on both iron-nickel alloy and mineral meteorite constituents is also surprisingly limited in the metallurgical, rock mechanics, and ceramics literature.

The mars sample return project

The Mars Sample Return (MSR) Project is underway. A 2003 mission to be launched on a Delta III Class vehicle and a 2005 mission launched on an Ariane 5 will culminate in carefully selected Mars samples arriving on Earth in 2008. NASA is the lead agency and will provide the Mars landed elements, namely, landers, rovers, and Mars ascent vehicles (MAVs). The French Space Agency CNES is the largest international partner and will provide for the joint NASA/CNES 2005 Mission the Ariane 5 launch and the Earth Return Mars Orbiter that will capture the sample canisters from the Mars parking orbits the MAVs place them in. The sample canisters will be returned to Earth aboard the CNES Orbiter in the Earth Entry Vehicles provided by NASA. Other national space agencies are also expected to participate in substantial roles. Italy is planning to provide a drill that will operate from the Landers to provide subsurface samples. Other experiments in addition to the MSR payload will also be carried on the Landers. This Papepwill present the current status of the design of the MSR missions and flight articles.

Aerodynamics of Stardust Sample Return Capsule

Successful return of interstellar dust and cometary material by the Stardust Sample Return Capsule requires an accurate description of the Earth entry vehicle''s aerodynamics. This desciption must span the hypersonic-rarefied, hypersonic-continuum, supersonic, transonic, and subsonic flow regimes. Data from numerous sources are compiled to accomplish this objective. These include Direct Simulation Monte Carlo analyses, thermochemical nonequilibrium computational fluid dynamics, transonic computational fluid dynamics, existing wind tunnel data, and new wind tunnel data. Four observations are highlighted: 1) a static instability is revealed in the free-molecular and early transitional-flow regime due to aft location of the vehicle''s center-of-gravity, 2) the aerodynamics across the hypersonic regime are compared with the Newtonian flow approximation and a correlation between the accuracy of the Newtonian flow assumption and the sonic line position is noted, 3) the primary effect of shape change due to ablation is shown to be a reduction in drag, and 4) a subsonic dynamic instability is revealed which will necessitate either a change in the vehicle''s center-of-gravity location or the use of a stabilizing drogue parachute.

Blunt body rarefied wakes for Earth entry

Direct simulation Monte Carlo (DSMC) and Navier-Stokes axisymmetric calculations are presented for hypersonic low-density flow about a 70-deg blunt cone afterbody configuration. The flow conditions simulated are those experienced by a space vehicle for an altitude range of 105-75 km during Earth entry. The entry velocity considered is 7 km/s. A steady vortex is predicted in the near wake for 75 and 85 km altitudes by both calculations. The flow remains attached for 95 and 105 km altitudes. Comparisons of DSMC and Navier-Stokes calculations for the wake flow field become less and less favorable with increasing altitude. Comparisons of surface quantities show reasonable agreement between DSMC and Navier-Stokes calculations along the forebody. However, the surface quantities along the afterbody calculated from DSMC and Navier-Stokes calculations differ significantly at the higher altitudes.

Effects of chemistry on blunt-body wake structure

Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Give a Fig

FICTION Give a Fig Shelba Cole Robison Ripping open the box lid, Leona plucked the cut glass from its tissue paper nest and signed deeply. The new cake plate was perfect for the contest. This year she needed a gimmick because Dorthus, her only serious baking competitor, was smart and getting smarter. Now all she had to do was slide her coconut marshmallow fluff atop the pedestal and it would look as fine as one of Queen Elizabeth's crowns. Dorthus really had an unfair advantage—why, she'd even studied home economics for two years at Lafayette Women's College. Not only did she cook well, she figured out clever angles—like her last year's contest entry, Seven-Up Cake. Pastor Beezer loved its catchy name, even if, in Leona's humble opinion, the cake turned out coarsely textured . Besides, Dorthus merely substituted Seven-Up for the milk in The Joy of Cooking's standard yellow cake recipe. Leona also found bottlecap decorations downright tacky. You couldn't even eat them. Leona's only asset was her raw talent and grit; fortunately, enough of both to capture every first place since the Isle of Hope Methodist women's missionary society contest began nine years ago. But now her talent and grit barely held their own against an almost professional home economist like Dorthus. Leona had to admit—last year her Chocolate and Red Earth entry barely nudged out that Seven-Up Cake. Katie Belle and Imodene told her she should be satisfied with her nine wins. "They've established your reputation . . . step back and let the young ones try their wings now." "Nine's puny. I want a whole decade," Leona told them. Ten wins would permit her to retire because, if the truth be known, she was plumb tired out from all the worrying and cooking. She wouldn't last much Shelba Cole Robison teaches creative writing at Saddleback College in Mission Viejo, California. She has compiled a novel-in-stories, from which we have published "Watch Over Me" (Vol. 17, No. 3, 1989), and "Burdens We Bear" (Vol. 19, No. 1, 1991). 45 longer. Ten wins would mean nobody, not even Dorthus, could equal her record during her lifetime. If Leona could win this year, she'd persuade Hazel, her best friend, to propose a Christian Cooks' Hall of Fame and to nominate Leona Coburn its first member. Public acknowledgment would be nice for a change. Despite her good works, not one person ever mentioned all that driving she did during World War II; nobody applauded the way she single-handedly raised, not only her daughter, but also Katie Belle's two, including ferrying them to piano lessons and recitals, chaperoning at dances and Wesleyan youth events. Only her friend, Hazel, ever understood the cross she had to bear with Imodene, who couldn't do anything unless Leona told her what and how to do it. She could stand a little recognition. Others grumbled that the contest was fixed because no one ever seemed to reach the finals except Leona and Dorthus. But it had never been easy for her—her cakes won because they looked and tasted the best. She worried about this year's entry though. The coconut marshmallow fluff looked pretty, to be sure, and tasted swell. She could just hear those little bitty pieces of coconut tease as if to say, "If you don't taste me you'll be sorry." But was it good enough to win? "It's a specially good cake, but there's nothing unusual about it," Leona bemoaned to Mildred of Mildred's Coiffures for Milady during her regular Saturday morning hair appointment. "I need something that nobody's heard of." "You know, Leona, I think I have just the thing for you." Mildred stopped in the middle of a pin curl. "You ever hear of Scripture Cake?" Leona thought for a moment. "Don't believe I have." "My nephew in Smyrna married a Pentecostal woman. I don't know much about her religion, but she cooks real good. When I visited them last summer she baked this wonderful cake—the ingredients come from Bible verses. With...